CN114611366A - Pyrolysis ablation simulation calculation method in spacecraft uncontrolled meteor reentry disintegration analysis - Google Patents
Pyrolysis ablation simulation calculation method in spacecraft uncontrolled meteor reentry disintegration analysis Download PDFInfo
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Abstract
The invention discloses a pyrolysis ablation simulation calculation method in uncontrolled airborne reentry disintegration analysis of a spacecraft, which comprises the following steps: step one, in the uncontrolled airborne reentry disintegration analysis of the spacecraft, establishing a reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm matched with the carbon-based composite material; solving the heat conduction inside the spacecraft and the disassembled shape of the spacecraft; and step three, solving the structure according to the heat conduction in the step two, and judging the failure sequence of the material, thereby obtaining the calculation method of the pyrolysis/ablation of the carbon-based composite material.
Description
Technical Field
The invention relates to the field of satellites and application products. More specifically, the invention relates to a pyrolysis ablation simulation calculation method used in spacecraft uncontrolled meteoron reentry disintegration analysis in the process of predicting and simulating uncontrolled spacecraft reentry atmosphere disintegration by a spacecraft measurement and control ground station.
Background
The analysis of the reentry of the spacecraft with the complex structure needs to firstly determine the criterion of material damage, namely the condition under which the material is damaged. For metal materials, a melting point temperature control model is adopted, and the surface of the metal material is molten and runs off under the action of pneumatic shearing force on the assumption that the surface temperature of the metal material reaches a melting point; for carbon-based composites, a pyrolysis/ablation control model is used, assuming that the structure is completely pyrolyzed or ablated, i.e., damaged. The method comprises the steps of carrying out aerodynamic heat/softening/melting or aerodynamic heat/pyrolysis/ablation computational analysis on a spacecraft along a small reentry angle uncontrolled trajectory, solving an ablation shape by applying a spring leaf stretching method to deform an unstructured moving grid technology, using tetrahedral grid outer boundary nodes for aerodynamic heat calculation, keeping grid unit and node number information unchanged in the ablation shape solving, only changing node coordinates, realizing data information transmission of different computational modules of the unstructured grid, carrying out coupling analysis on the ablation shape moving grid, three-dimensional finite element heat transmission and aerodynamic heat, and giving theoretical prediction on the disassembly condition of the whole spacecraft and parts thereof.
The existing simulation method mainly comprises three steps:
step one, in the uncontrolled airborne reentry disintegration analysis of the spacecraft, establishing a reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm matched with the carbon-based composite material;
solving the heat conduction inside the spacecraft and the disassembled shape of the spacecraft;
and step three, solving the structure according to the heat conduction in the step three, judging the failure sequence of the material, and further obtaining the calculation method of the pyrolysis/ablation of the carbon-based composite material.
Disclosure of Invention
An object of the present invention is to solve at least the above problems and/or disadvantages and to provide at least the advantages described hereinafter.
To achieve these objects and other advantages in accordance with the purpose of the invention, there is provided a method for pyrolytic ablation simulation calculation in uncontrolled airborne reentry disintegration analysis of a spacecraft, comprising:
step one, in uncontrolled meteor reentry disintegration analysis of an aircraft, establishing a reentry pneumatic thermal ablation and internal temperature distribution fast algorithm matched with a carbon-based composite material;
solving the heat conduction inside the spacecraft and the disassembled shape of the spacecraft;
and step three, solving the structure according to the heat conduction in the step two, and judging the failure sequence of the material, thereby obtaining the calculation method of the pyrolysis/ablation of the carbon-based composite material.
Preferably, in step one, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to include:
s10, establishing a finite element algorithm suitable for the aerodynamic thermal environment structure to respond to the transient temperature field, and further forming a preliminary spacecraft disintegration criterion model;
and S11, for the carbon-based composite material, adopting a pyrolysis-ablation control model, assuming that the structure is completely pyrolyzed or ablated, and adopting the established spacecraft disintegration criterion model to carry out coupling solution on the aerodynamic heat and the structure thermal response.
Preferably, in S10, the establishing of the finite element algorithm is configured to include:
s101, solving a spacecraft surface structure temperature field by adopting a tetrahedral four-node finite element method based on a three-dimensional transient heat conduction equation, and deriving a finite element method structure thermal response balance equation by using functional variational components;
s102, based on finite element method overall synthesis, the rigidity matrix has regularity of symmetry positive definite, high sparseness and non-zero element distribution, a one-dimensional variable broadband storage technology is developed, and a numerical solving strategy of a three-dimensional finite element heat transfer calculation model is constructed;
s103, constructing a two-point backward difference format, a Crank-Nicolson format and a Galerkin format for solving the three-dimensional transient temperature field based on a method of the centralized variable temperature matrix coefficient, and further obtaining a finite element algorithm suitable for the aerodynamic thermal environment structure to respond to the transient temperature field.
Preferably, in the second step, the problem of the three-dimensional temperature field is solved by adopting a tetrahedral four-node unit based on a three-dimensional heat conduction equation according to the ablation condition obtained in the first step, so that the heat conduction condition in the reentry process of the spacecraft is simulated;
in the solving process of the three-dimensional temperature field, a Kriging spatial interpolation algorithm is adopted to obtain global three-dimensional data by using partial three-dimensional data.
Preferably, in step three, the failure sequence of the material is:
for the metal material, a melting point temperature control model is adopted, and the surface temperature of the metal material is assumed to reach the melting point and is lost under the action of pneumatic shearing force;
for carbon-based composites, a pyrolysis/ablation control model is employed, assuming that the structure is fully pyrolyzed or ablated.
The invention at least comprises the following beneficial effects: in the simulated solving process, a Kriging spatial interpolation algorithm is introduced in the step, and as a statistical method for spatial interpolation, the global three-dimensional data can be obtained by utilizing partial three-dimensional data, so that the calculated amount in the process of solving the three-dimensional temperature field is reduced, and the calculation time is shortened.
Additional advantages, objects, and features of the invention will be set forth in part in the description which follows and in part will become apparent to those having ordinary skill in the art upon examination of the following or may be learned from practice of the invention.
Drawings
FIG. 1 is a graphical illustration of a reentry trajectory of a gas cylinder;
FIG. 2 is a graph of ablation and temperature profile of a cylinder at 85km height;
figure 3 is an ablation and temperature profile of the cylinder at 83.3km height.
Detailed Description
The present invention is described in further detail below to enable those skilled in the art to practice the invention with reference to the description.
The invention relates to an implementation form of a pyrolysis ablation simulation calculation method in spacecraft uncontrolled meteor reentry disintegration analysis, which comprises the following steps:
step one, in the uncontrolled airborne reentry disintegration analysis of the spacecraft, establishing a reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm matched with the carbon-based composite material;
solving the heat conduction inside the spacecraft and the disassembled shape of the spacecraft;
and step three, solving the structure according to the heat conduction in the step two, judging the failure sequence of the material, and further obtaining the calculation method for the pyrolysis/ablation of the carbon-based composite material.
In the carbon-based composite reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm of step one, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to include:
the method mainly simulates the internal temperature change of the carbon-based composite material due to the pneumatic action in the reentry process of the spacecraft, and then simulates the thermal ablation condition of the material.
Starting from a three-dimensional transient heat conduction equation, a tetrahedral four-node finite element method is provided to solve a spacecraft surface structure temperature field, and a finite element method structure thermal response balance equation is derived by using functional variational components. Constructing a numerical solving strategy of a three-dimensional finite element heat transfer calculation model, utilizing a finite element method to carry out overall synthesis to ensure that the rigidity matrix has regularity of symmetrical positive definite, high sparsity and non-zero element distribution, developing a one-dimensional variable broadband storage technology and solving the data storage problem of a large sparse matrix; the method of the centralized variable temperature matrix coefficient is applied to effectively inhibit the temperature oscillation phenomenon in the solving process, construct a two-point backward difference format, a Crank-Nicolson format and a Galerkin format for solving the three-dimensional transient temperature field, and establish a finite element algorithm suitable for the pneumatic thermal environment structure to respond to the transient temperature field. A spacecraft disintegration criterion model with a complex structure is preliminarily formed, and a pyrolysis/ablation control model is adopted for the carbon-based composite material, and the structure is supposed to be completely pyrolyzed or ablated to be damaged. The established model is applied to solve the coupling of aerodynamic heat and structural thermal response, the transient aerodynamic thermal environment and material disintegration damage of the two cabin bodies of the air ship similar to the skater ship I and the thin-shell structure which fall along the small reentry angle are calculated and analyzed, the result is well matched with the finite element software calculation result, and the reliability of the heat transfer model is verified.
Research shows that the ablation process of the carbon-based material can be described by a layered model, and the part of the material inside, which is close to the heating surface, is a carbonized layer which is a porous material formed after the original material is completely pyrolyzed. The pyrolysis zone in which the pyrolysis reaction is generated is arranged between the carbonization layer and the original material layer, the pyrolysis reaction can absorb heat, and pyrolysis gas generated by the reaction can flow in the carbonization layer to play a role in blocking heat transfer. Below the pyrolysis zone is the starting material that has not undergone pyrolysis.
In another example, in the second step, according to the ablation condition obtained in the first step, a three-dimensional heat conduction equation is utilized, and a tetrahedral four-node unit is adopted to solve a three-dimensional temperature field problem, so that the heat conduction condition in the reentry process of the spacecraft is simulated.
In the solving process, a Kriging spatial interpolation algorithm is introduced in the step, and as a statistical method for spatial interpolation, global three-dimensional data can be obtained by utilizing partial three-dimensional data, so that the calculated amount in the process of solving the three-dimensional temperature field is reduced, and the calculating time is shortened.
In order to verify the ablation shape solving technology and the moving boundary aerodynamic heat/ablation and internal heat conduction coupling solving technology of the complex structure of the spacecraft, a gas cylinder component is taken as an example, the typical high ablation and surface temperature distribution of a gas cylinder of a multi-time disintegration component at an attack angle of 20 degrees are calculated, and the gas cylinder component is made of an aluminum alloy material which is easy to ablate in order to observe the effect of grid deformation. Fig. 1 is a reentry ballistic curve of the gas cylinder, the initial ballistic height is 90km, the initial reentry speed is 7.48km/s, the flying attack angle of the gas cylinder is assumed to be 20 degrees, fig. 2-3 are the results of calculation of ballistic ablation and temperature distribution of the gas cylinder along the wind, the temperature of the gas cylinder on the windward side is higher, and ablation is more obvious. In the case of flight at an angle of attack, the aircraft will be ablated into an asymmetric profile, with disintegration occurring first in the face region of the wind.
Through the test, the method added with the Kriging spatial interpolation algorithm and the original method obtain the result with the error within 1 percent, but due to the introduction of the Kriging spatial interpolation algorithm, the second step of the method reduces the calculation amount by 80 percent and shortens the calculation time by nearly 30 percent, so the method has obvious advantages in the aspect of calculation efficiency compared with the original method.
In the pyrolysis/ablation calculation method of the carbon-based composite material in the third step, the damage analysis of the spacecraft with the complex structure needs to firstly determine the criterion of material damage, namely, the condition under which the material fails. The failure sequence of the material is judged according to the heat conduction simulation result of the previous step.
For metal (alloy) materials, a melting point temperature control model is adopted, and the surface of the metal material is molten and runs off under the action of pneumatic shearing force on the assumption that the surface temperature of the metal material reaches a melting point; for carbon-based composites, a pyrolytic/ablative control model is used, assuming that the structure is completely pyrolytic or ablated, i.e., damaged. Based on surface aerodynamic thermal data of the low-orbit spacecraft with the service expiration at the height of 120km to 90km, the theoretical prediction of the disintegration condition of the low-orbit spacecraft thin shell structure and parts thereof is given by carrying out aerodynamic thermal/fusion or aerodynamic thermal/pyrolysis/ablation calculation analysis along the small reentry angle uncontrolled trajectory, and data support is provided for the drop point analysis of the debris.
For metal materials, a melting point temperature control model is adopted, and the surface of the metal material is molten and runs off under the action of pneumatic shearing force on the assumption that the surface temperature of the metal material reaches a melting point; for carbon-based composites, a pyrolytic/ablative control model is used, assuming that the structure is completely pyrolytic or ablated, i.e., damaged. The method comprises the steps of carrying out aerodynamic heat/softening/melting or aerodynamic heat/pyrolysis/ablation computational analysis on a spacecraft along a small reentry angle uncontrolled trajectory, solving an ablation shape by applying a spring leaf stretching method to deform an unstructured moving grid technology, using tetrahedral grid outer boundary nodes for aerodynamic heat calculation, keeping grid unit and node number information unchanged in the ablation shape solving, only changing node coordinates, realizing data information transmission of different computational modules of the unstructured grid, carrying out coupling analysis on the ablation shape moving grid, three-dimensional finite element heat transmission and aerodynamic heat, and giving theoretical prediction on the disassembly condition of the whole spacecraft and parts thereof.
The above scheme is merely illustrative of a preferred example, and is not limiting. In the implementation of the invention, appropriate replacement and/or modification can be carried out according to the requirements of users.
The number of apparatuses and the scale of the process described herein are intended to simplify the description of the present invention. Applications, modifications and variations of the present invention will be apparent to those skilled in the art.
While embodiments of the invention have been disclosed above, it is not intended to be limited to the uses set forth in the specification and examples. It can be applied to all kinds of fields suitable for the present invention. Additional modifications will readily occur to those skilled in the art. The invention is therefore not to be limited to the specific details described herein, without departing from the general concept as defined by the appended claims and their equivalents.
Claims (4)
1. A pyrolysis ablation simulation calculation method in spacecraft uncontrolled meteor reentry disintegration analysis is characterized by comprising the following steps:
step one, in the uncontrolled airborne reentry disintegration analysis of the spacecraft, establishing a reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm matched with the carbon-based composite material;
solving the heat conduction inside the spacecraft and the disassembled shape of the spacecraft;
step three, solving the structure according to the heat conduction in the step two, and judging the failure sequence of the material, thereby obtaining a calculation method of the pyrolysis/ablation of the carbon-based composite material;
in the second step, solving a three-dimensional temperature field problem by adopting a tetrahedral four-node unit based on a three-dimensional heat conduction equation according to the ablation condition obtained in the first step, thereby simulating the heat conduction condition in the reentry process of the spacecraft;
in the solving process of the three-dimensional temperature field, a Kriging spatial interpolation algorithm is adopted to obtain global three-dimensional data by utilizing partial three-dimensional data.
2. The method for calculation of simulation of pyrolytic ablation in uncontrolled airborne reentry disassembly analysis of spacecraft of claim 1, wherein in step one, the reentry aerodynamic thermal ablation and internal temperature distribution fast algorithm is configured to comprise:
s10, establishing a finite element algorithm suitable for the aerodynamic thermal environment structure to respond to the transient temperature field, and further forming a preliminary spacecraft disintegration criterion model;
and S11, for the carbon-based composite material, adopting a pyrolysis-ablation control model, assuming that the structure is completely pyrolyzed or ablated, and adopting the established spacecraft disintegration criterion model to carry out coupling solution on the aerodynamic heat and the structure thermal response.
3. The method for pyrolytic ablation simulation calculation in an uncontrolled merle reentry disassembly analysis of spacecraft of claim 2, wherein in S10, the establishing of finite element algorithm is configured to include:
s101, solving a spacecraft surface structure temperature field by adopting a tetrahedral four-node finite element method based on a three-dimensional transient heat conduction equation, and deriving a finite element method structure thermal response balance equation by using functional variational components;
s102, based on finite element method overall synthesis, the rigidity matrix has regularity of symmetry positive definite, high sparseness and non-zero element distribution, a one-dimensional variable broadband storage technology is developed, and a numerical solving strategy of a three-dimensional finite element heat transfer calculation model is constructed;
s103, constructing a two-point backward difference format, a Crank-Nicolson format and a Galerkin format for solving the three-dimensional transient temperature field based on a method of the centralized variable temperature matrix coefficient, and further obtaining a finite element algorithm suitable for the aerodynamic thermal environment structure to respond to the transient temperature field.
4. The method for pyrolytic ablation simulation calculation in uncontrolled airborne reentry disintegration analysis of spacecraft of claim 1, wherein in step three, the failure order of the material is:
for the metal material, a melting point temperature control model is adopted, and the surface temperature of the metal material is supposed to reach the melting point and run off under the action of pneumatic shearing force;
for carbon-based composites, a pyrolysis/ablation control model is employed, assuming that the structure is fully pyrolyzed or ablated.
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| CN117174216A (en) * | 2023-10-24 | 2023-12-05 | 浙江大学 | Laminated composite thermal response analysis method, electronic device, and readable storage medium |
| CN117174216B (en) * | 2023-10-24 | 2024-02-06 | 浙江大学 | Laminated composite thermal response analysis method, electronic device, and readable storage medium |
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