CN111241634B - Analysis and forecast method for reentry of spacecraft into meteor space - Google Patents

Abstract

The invention discloses an analysis and forecast method for reentry of spacecraft into meteor space, which comprises the following steps: (1) analyzing the structure of the spacecraft and constructing a disintegration three-level model; (2) performing aerodynamic, and flight characteristics computational analysis on the system/subsystem level and the component level; (3) determining a system/subsystem level disintegration criterion parameter and a condition boundary parameter; (4) calculating and simulating a system/subsystem level flight path; (5) performing calculation simulation analysis on the system/subsystem hierarchical structure disintegration damage; (6) determining a component level disintegration criterion parameter and a condition boundary parameter; (7) calculating and simulating a component level flight path; (8) performing calculation and simulation analysis on component hierarchical structure disintegration; (9) constructing a piece strip model of a fragment/particle level; (10) analyzing the survivability of the spacecraft fragments and the boundary of a simulation calculation condition; (11) calculating and simulating the fragment/particle level aerodynamic force and flight path; (12) analyzing and evaluating the falling area of the fragment/particle level; (13) and (4) ground risk assessment.

Description

Analysis and forecast method for reentry of spacecraft into meteor space

Technical Field

The invention relates to an analysis and forecast method for reentry of spacecraft into meteorology, aiming at the description of physical research objects of various components or fragments possibly appearing in the process of reentry of spacecraft into meteorology in the aspect of geometric physics, and carrying out aerodynamic force/aerodynamic heat/disintegration analysis, survival of debris and ground risk assessment of various research objects in the flight process.

Background

With the increasing space activities outside the atmosphere, the problem of reentry of spacecrafts into space is also gradually concerned and paid attention. The process of spacecraft meteorology disintegration refers in particular to the process of enabling various invalid satellites, non-recovery type artificial celestial bodies (such as an orbit cabin and a space station of a manned spacecraft) which run in orbit, rocket boosters or carrier wreckages and the like to enter the earth atmosphere to the ground after uncontrolled or controlled orbital transfer is finished. These types of flyers are often not designed for flying in the atmosphere and therefore do not have a specific aerodynamic layout profile and thermal protection structure. After the aerolite flies into the atmosphere again at an ultrahigh speed, under the action of strong aerodynamic force/aerodynamic heat, the original configuration of the aerolite is disintegrated, the metal material is softened and melted, the composite material is pyrolyzed/ablated, and violent phenomena such as combustion of certain materials, explosion or explosion of container parts and the like are caused during the aerolite process. The understanding and understanding of the reentry and meteorology process of a spacecraft is a prerequisite for modeling and analysis of forecasts.

Spacecraft merle is typically a non-conventional reentry problem (conventional reentry: recovery or aggressor; non-conventional reentry: bump or merle). The main technical problems in the process of aerolisation of the spacecraft comprise four aspects of flight motion, aerodynamic force, aerodynamic heat and structural disintegration. In the last two thirty years, some research works aiming at spacecraft and space debris reentering meteor space have been carried out at home and abroad, but due to the complexity of problems caused by multidisciplinary crossing, the related research and application level is still very limited, and some basic problems are not well solved.

Various phenomena in the process of spacecrafts merning have great randomness, and from the perspective of solving engineering problems, the uncertainty of a system cannot be necessarily reflected by carrying out complex spacecraft fine modeling under a limited condition state. Therefore, the model and the method provided by the invention integrate the characteristics of a spacecraft-oriented method and an object-oriented method. The three-layer model and the parameter statistical method based on the condition boundary can cover the structural composition and various geometrical and physical characteristics of fragments of a spacecraft disintegration research object, aerodynamic force, aerodynamic heat and flight motion characteristics of the research object can be calculated and analyzed based on the geometrical shape and the physical attribute characteristics, and the method has universal adaptability to different spacecrafts or different disintegration conditions through application of adjusting parameters in the method.

Disclosure of Invention

The invention aims to provide an analysis and forecast method for reentry and merry of a spacecraft, which carries out modeling simulation on aerodynamic force, aerodynamic heat and disintegration damage in the process of reentry and merry flight of the spacecraft, determines description of physical research objects of various components or fragments possibly occurring or generated in the process of disintegrating the spacecraft in the aspect of geometric physics through a disintegration model of the method, and can simulate the number, size, quality and falling area range of the components or fragment debris in the process of disintegrating the spacecraft through the parameter statistical analysis set by condition boundaries of the method so as to further carry out analysis and evaluation on ground risks.

To achieve these objects and other advantages in accordance with the purpose of the invention, there is provided an analytical forecasting method for spacecraft reentry meteority, comprising the steps of:

analyzing a spacecraft structure and constructing a disintegration three-level model; the disassembly three-level model comprises a spacecraft system/subsystem level, a spacecraft component level, and a spacecraft debris/particle level; calculating and analyzing aerodynamic force, aerodynamic heat and flight characteristics of a spacecraft system/subsystem level and a spacecraft component level; step three, determining a spacecraft system/subsystem level disintegration criterion parameter and a condition boundary parameter; fourthly, calculating and simulating the flight path of the spacecraft system/subsystem level; fifthly, decomposing, destroying, calculating, simulating and analyzing the hierarchical structure of the spacecraft system/subsystem; sixthly, determining a spacecraft component level disintegration criterion parameter and a condition boundary parameter; seventhly, calculating and simulating a flight path of the spacecraft component level; eighthly, calculating, simulating and analyzing the decomposition and damage of the hierarchical structure of the spacecraft component; constructing a spacecraft fragment/particle level block strip model; tenthly, analyzing the survivability of the spacecraft fragments and the boundary of a simulation calculation condition; step eleven, calculating and simulating the aerodynamic force and flight path of the spacecraft fragment/particle level; twelfth, analyzing and evaluating a falling area of the spacecraft fragment/particle level; and thirteen, ground risk assessment.

Preferably, the spacecraft system in the spacecraft system/subsystem hierarchy is in the form of a whole spacecraft, and the spacecraft subsystem is generally a relatively complete section of an oversized spacecraft or a complete cabin section with external accessory parts falling off; the spacecraft component level refers to an external object or an internal object which has certain functionality and structure shape certainty inside and outside a spacecraft cabin; the spacecraft debris/particle level refers to the residue of the original geometric form which cannot be visually identified, and belongs to an abstract model.

Preferably, in the second step, the aerodynamic force calculation and analysis adopts an engineering calculation method or a numerical simulation method; based on the characteristics of the meteorolite and the complex shape of the wreckage of the spacecraft, the aerodynamic force calculation is mainly based on an engineering method; performing supplement and comparison and verification of typical state calculation results by using numerical simulation;

in the second step, the pneumatic thermal parameters before the whole system/subsystem level spacecraft is disassembled are obtained by numerical simulation such as a DSMC method; from the engineering application angle of spacecraft meteor analysis and forecast, a rapid engineering calculation method is adopted along a large number of thermal environment parameters of a trajectory; due to the inexact expectation of the shape of the deformed falling body and the disassembled part, the local details are properly simplified in the calculation of the thermal environment engineering.

Preferably, the aerodynamic engineering calculation method adopts a surface element method, and comprises Newton theory and correction of continuous flow region, molecular collision theory of free molecular flow region, and overlapping formula processing of thin transition flow region; the numerical simulation method comprises any one of Boltzmann model equation unified algorithm, DSMC numerical simulation, N-S/DSMC coupling algorithm and NS equation numerical simulation;

in the specific processing of the pneumatic thermal parameter acquisition, the meteor windward side is subjected to axisymmetric matching assumption, and the problem of streaming with an attack angle and a sideslip angle is converted into the problem of simple total attack angle streaming by adopting an equivalent method; the problem of zero-attack-angle streaming can be further solved by applying an equivalent spherical cone method. After the transformation, the aerodynamic heat problems of an attack angle and a belt sideslip can be calculated by using a zero-attack-angle spherical cone aerodynamic heat calculation method.

Preferably, in the second step, the calculation and analysis of the flight characteristics adopt a six-degree-of-freedom or three-degree-of-freedom trajectory calculation method; when the movement details before disassembly need to be mastered, a six-degree-of-freedom ballistic equation can be adopted; the ballistic equation set can be solved step by using a 4-order Runge-Kutta method integral.

Preferably, in the third step, aiming at flight environment, aerodynamic force/heat, flight motion and attitude parameters with uncertainty in simulation modeling analysis of the meteor process of the spacecraft, the upper limit and the lower limit, namely the conditional boundary, of the flight environment, aerodynamic force/heat, flight motion and attitude parameters are evaluated and determined according to the existing basic research result; constructing a proper parameter distribution mode in the boundary value domains of the conditions during analysis of the meteority process; constructing a disintegration condition criterion of structure disintegration damage; based on a statistical analysis correlation mathematical method, combining flight motion-force-heat-structure destruction disintegration calculation to carry out quantitative analysis on the concerned target parameters; obtaining statistically significant spacecraft flight path, disintegration process, debris viability and ground risk analysis evaluation results in the form of distribution bands with a certain degree of confidence; the parameter distribution refers to each condition parameter with uncertainty, the condition parameters are set to be normal distribution with a reference value as a center, namely Gaussian distribution, the random value range is based on the reference value, the deviation is plus-minus three times of standard deviation, namely the parameter coverage rate is about 99.5% theoretically;

the condition state parameters related to the condition boundary cover performance parameters in the aspects of flight motion, aerodynamic force and aerodynamic heat, and comprise key time points such as reentry time parameters and disintegration time parameters, and physical parameters of a penetration process such as atmospheric density, aerodynamic force coefficients and the like; in the preliminary induction, the statistical analysis parameters related to the spacecraft merle analysis forecasting method comprise: initial position, initial velocity, aerodynamic coefficient, atmospheric density, height of disintegration, property of disintegration fragments, and attitude of disintegration fragments;

the disintegration condition criteria include: height criteria: i.e. a melting criterion is set for a certain height value for some form of disintegration process to occur: namely, the metal material is set to have a certain form of disintegration process at the melting temperature; and (3) pyrolysis criterion: namely, the composite material is set to be subjected to a certain form of disintegration process after being completely pyrolyzed at the pyrolysis temperature; temperature criterion: namely, the object plane temperature is considered to be up to a certain set value and a certain form of disintegration process will occur; heat flow criteria: that is, it is considered that some form of disintegration process will occur when the accumulated heat flow or the instantaneous heat flow reaches a certain set value; dynamic pressure criterion: namely, the dynamic pressure reaches a certain set value and a certain form of disintegration process is generated; and (3) comprehensive criteria: the combination of the two or more criteria can include the weight sum, the achievement of any criterion, the achievement of all criteria and the like; the melting criterion and the pyrolysis criterion are a composite form of a temperature criterion and a heat flow criterion; the melting rule is listed to facilitate the application analysis of the metal material with the largest proportion in the current spacecraft structural materials; the pyrolysis criterion is listed to carry out application analysis on the main material composite materials of the container class with special requirements on the current spacecraft.

Preferably, in the fifth step, the analysis result of aerodynamic force and aerodynamic heat is used as the boundary condition of the object plane of the structural stress analysis; obtaining physical property parameters changing along with temperature and thermal stress caused by temperature difference by thermal analysis in the structure; under the comprehensive action of object surface load transfer stress and thermal stress, analyzing structural damage according to the total stress distribution condition of the material and parameters such as Young modulus, Poisson ratio and the like under corresponding conditions;

in the sixth step, the determination of the spacecraft component level disintegration criterion parameter and the conditional boundary parameter is consistent with the determination of the spacecraft system/subsystem level disintegration criterion parameter and the conditional boundary parameter in the third step, and the parameters include: the initial position and the initial speed of the component level and the parameters of the inherited system/subsystem level disintegration time; the rest of the components comprise aerodynamic coefficient, atmospheric density and disintegration height;

in the seventh step, a three-degree-of-freedom ballistic equation and a aerodynamic heat rapid calculation method are adopted for calculating and simulating the component-level flight path; and calculating attenuation change of the mass of the research object according to the fusion ablation rate in the simulation process.

Preferably, in the ninth step, in the modeling of the shapes of the spacecraft fragment/particle levels, the geometric features of the spacecraft fragment/particle levels are highly abstracted; the geometric modeling abstraction is based on the principle of a complete induction method, and different geometric characteristics can be covered, so that corresponding pneumatic characteristics and flight motion characteristics are covered and embodied; the specific method is that the fragments/particles are divided into the following parts according to geometrical characteristics: blocks, strips, sheets; this mode of division of the spacecraft debris/particle level is therefore referred to as the patch model.

Preferably, in the eleventh step, aerodynamic force and flight path of the block strip model are calculated and simulated according to the block strip model, and the aerodynamic force adopts aerodynamic force coefficient tables with different shapes for the block strip to perform interpolation; and calculating and simulating the flight path by adopting a three-degree-of-freedom trajectory equation.

Preferably, in the twelfth step, the landing area scattering is analyzed and evaluated according to the track data of the debris fragments obtained in the eleventh step; the landing zone range depends on the following factors: flight parameters and uncertainty at the time of disassembling the whole device and the parts, and the material, shape, size and distribution rule of debris fragments.

The invention at least comprises the following beneficial effects: the invention mainly aims at the meteorite process of the unconventional reentry (namely, non-recovery and non-attack, mainly meteorite) spacecraft, establishes a hierarchical mode for describing the geometric configuration of a research object at different disintegration stages and a distribution form such as scale quality of the hierarchical mode, provides a reasonable receptor object for aerodynamic force/aerodynamic heat calculation analysis of the research object in the flight process, provides a basis for flight motion calculation of the research object in the disintegration process, and performs motion-force-heat-structure damage disintegration calculation on the disintegration physical process by combining a mathematical method based on a statistical theory, thereby analyzing and forecasting the reentry meteorite flight track, the disintegration process, the survival of debris and the ground risk of the spacecraft. (1) The three-level model has clear and definite structure definition of a research object, is refined and applicable and is convenient for engineering operation. (2) The definition of the spacecraft component hierarchy facilitates the exploration of the details of the merle disintegration process. (3) The "patch-strip model" of the spacecraft chip/particle level can represent different geometrical characteristics and corresponding aerodynamic/thermal and flight motion characteristics. (4) The three-level model can meet the requirement of engineering analysis and is also beneficial to digging basic scientific problems to promote the deep research of disciplinary major.

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.

Description of the drawings:

FIG. 1 is a flow chart of the method for analyzing and forecasting reentry of a spacecraft of the present invention;

FIG. 2 is a system/subsystem hierarchy of a spacecraft integer-reentrant merle-disassembly three-level model;

FIG. 3 is a component hierarchy of a typical component-reentrant merle three-level model of a spacecraft;

FIG. 4 shows that six-component parameters of typical aerodynamic performance of a whole spacecraft re-entering flight change with an attack angle;

FIG. 5 is a graphical representation of the results of numerical simulations of an entire spacecraft and typical components;

FIG. 6 is the result of an aerodynamic heating calculation under typical conditions at the system/subsystem level;

FIG. 7 is the result of an aerodynamic thermal calculation of a part-level study;

FIG. 8 is a system/subsystem flight motion calculation for two different initial condition states;

FIG. 9 is a cloud of surface temperature distributions under typical conditions at the system/subsystem level;

FIG. 10 is a system/subsystem level, partial typical part flight path calculation and relationship;

FIG. 11 is a graph of surface temperature distributions at different heights for a typical component at a component level;

FIG. 12 is a timing diagram of a spacecraft reentry merle disruption;

FIG. 13 is a calculation result of a drop point range where a fragment may be generated under a certain set condition;

FIG. 14 is a schematic view of the position of the subgasket position and the grid void safety region in the last 138 rounds of the spacecraft to be meterled;

FIG. 15 is a schematic diagram of the last 5 rounds of trajectory and at least one half of the safe region range of the spacecraft to be meterled;

FIG. 16 is a graph showing the probability of hitting a remaining debris/fragment on a unit area of a time table at an inclination angle of 43.8 degrees;

FIG. 17 is a schematic illustration of three levels of a spacecraft merle disintegration three-level model.

The specific implementation mode is as follows:

the present invention is further described in detail below with reference to the attached drawings so that those skilled in the art can implement the invention by referring to the description text.

It will be understood that terms such as "having," "including," and "comprising," as used herein, do not preclude the presence or addition of one or more other elements or groups thereof.

The invention provides an analysis and forecast method for reentry of spacecraft into meteor space, which comprises the following steps:

analyzing a spacecraft structure and constructing a disintegration three-level model;

the method comprises the following steps of providing a novel and simple spacecraft merle disassembly model, namely a three-level model, and providing a necessary object geometry premise for analyzing and forecasting the merle process of the spacecraft; the basic idea of the model is an object-oriented method, and simultaneously, the thinking oriented to a spacecraft method is integrated; the three levels of the spacecraft merle three-level model (see FIG. 17) are: a system/subsystem; a component II; (iii) chips/particles;

for a spacecraft to be analyzed and evaluated, which is about to fall off, the overall appearance and structure of the spacecraft are analyzed so as to lay a foundation for constructing a disintegration research object; these structural analyses include externally apparent solar panels, antennas, etc.; also comprises the composition of internal components and the assembly mode thereof, and the internal components comprise containers, optical instruments, storage batteries and the like; besides obtaining the overall spacecraft and the geometric parameters of the main parts of the spacecraft, the relevant quality and physical parameters are also indispensable; according to the analysis condition of the spacecraft structure, the construction of a disintegration three-level model can be carried out by combining the estimated reentry merle flight trajectory form;

the system/subsystem level of the three-level model of the spaceborne meteorology disintegration, namely the form of the whole spacecraft; for an original on-orbit spacecraft, it may generally be in the form of a solar panel or not; for an oversize in-orbit operation spacecraft such as an international space station, the oversize in-orbit operation spacecraft can be analyzed by being divided into a plurality of subsystems in some merle stages; through example analysis (see the following specific embodiment 1), it can be known that the accuracy of state parameters, especially altitude, speed, trajectory inclination angle and flight azimuth angle, at the time of spacecraft disintegration plays a very decisive role in the estimation of the landing area range of subsequent survival debris, and also has an influence on the debris viability analysis and the ground risk assessment; this is the most significant meaning of specifying the system/subsystem hierarchy in a three-level model;

the component level of the three-level model of the spacecraft merle disintegration refers to an external object or a content with certain functionality and structural shape certainty inside and outside a spacecraft cabin, and comprises various functional components, containers and the like. When the falling space reaches a certain height, the shell and the attached connecting and fixing structure of the whole spacecraft are invalid under the action of aerodynamic force/heat, and various parts in the spacecraft are scattered under the impact of airflow; the disassembly time parameter of the whole device is the initial parameter of each part bearing severe aerodynamic force/heat and motion analysis;

the debris/particle level of the three-level model of the spacecrafts merle disintegration refers to that the residues of the original geometric morphology cannot be visually identified and belongs to an abstract model; the debris/particle source has several aspects, namely, the residue after the fusion ablation of the whole spacecraft or the part, the debris generated by the tearing of the whole spacecraft or the part under the shearing action of the airflow, and the residue after the recondensation and survival of the fusion; setting the upper and lower limits of the characteristic scale of the fragments/particles according to observation results and statistical analysis experience; the technical analysis of the fragments/particles no longer takes into account the aerodynamic thermal effects, i.e. they are treated as falling objects; the total amount of debris/particles seals the mass conservation and material properties of the merle spacecraft, and simultaneously the part deducting combustion or evolving into ultrafine dust particles is considered;

calculating and analyzing aerodynamic force, aerodynamic heat and flight characteristics of a spacecraft system/subsystem level and a spacecraft component level; the method comprises the following steps:

step 2.1, aerodynamic force evaluation:

the aerodynamic force calculation method can pertinently select an engineering calculation method or a numerical simulation method according to different research objects or the requirements of technical indexes; based on the characteristics of the complex shape of the meteorolite and the carcasses thereof, the aerodynamic force calculation is mainly based on an engineering method; performing supplement and comparison and verification of typical state calculation results by using numerical simulation; in meteor analysis forecast modeling, aerodynamic calculation analysis provides aerodynamic force and aerodynamic moment data of meteor in the flight process; at this time, aerodynamic data mainly provide results related to aircraft drag characteristics and static stability;

under the condition of hypersonic speed, the molecular collision theory of free molecular flow and the Newton theory of continuous flow are mature aerodynamic force rapid calculation methods; a plurality of optional bridging formulas are arranged in the transition flow area; one point to be concerned is that if partial remains are close to the earth surface and the speed is reduced below the super-sonic speed, or a numerical simulation method is adopted to obtain the pneumatic characteristics of the remains; experience gained from preliminary evaluations shows that debris fragments from such lower speed flights are not destroyed by aerodynamic thermal effects and it is important that their local ballistic inclination is generally close to-90 degrees (close to vertical drop), i.e. their drop point does not change significantly after flight through this type of mode.

In conclusion, the aerodynamic force calculation of the spacecraft in the falling flight process can be realized by adopting rapid engineering calculation; during specific implementation, the aerodynamic force can be interpolated by using a data table or directly coupled with the aerodynamic force for fast calculation;

step 2.2, aerodynamic heat assessment:

in a typical state, the pneumatic thermal parameters before the whole system/subsystem level spacecraft is disassembled are obtained by numerical simulation such as a DSMC method; from the engineering application angle of spacecraft meteor analysis and forecast, a rapid engineering calculation method is adopted along a large number of thermal environment parameters of a trajectory; due to the inexact expectation of the shapes of the deformed falling bodies and the disassembled parts, the local details are properly simplified in the thermal environment engineering calculation; in the specific treatment, the windward side of the meteorolite is subjected to an axisymmetric analogy hypothesis, and the problem of the streaming with an attack angle and a sideslip angle is converted into the problem of the simple total attack angle streaming by adopting an equivalent method; the problem of zero-attack-angle streaming can be further solved by applying an equivalent spherical cone method; after the transformation, the aerodynamic heat problems of an attack angle and a belt sideslip can be calculated by using a zero-attack-angle spherical cone aerodynamic heat calculation method;

by confirming and processing the characteristic size and the characteristic thickness of the spacecraft system/subsystem and component level, even a zero-dimensional model can be adopted to carry out more rapid estimation and analysis on the aerodynamic heat;

step 2.3, flight motion evaluation:

the meteor body flying motion solving method can adopt a mature six-degree-of-freedom or three-degree-of-freedom trajectory calculation method in principle; when the movement details before disassembly need to be mastered, a six-degree-of-freedom ballistic equation can be adopted; the ballistic equation set can be solved step by adopting a 4-order Runge-Kutta method;

based on the characteristics of complex shape change of the meteorolite integer device and the components in the meteorolite disintegration process, the quality characteristic is also the quantity which is difficult to grasp for the flyers; when the rotational inertia is uncertain, the three-degree-of-freedom trajectory equation is adopted to solve the trajectory of the meteorolite and the debris thereof; estimating one or more possible corresponding attitude angles of the static stable points according to the specific analysis of the aerodynamic characteristics of the meteorole in the motion attitude, and taking the static stable points as the basis of an attitude angle parameter distribution statistical model so as to carry out statistical quantification on the meteorole aerodynamic force; based on the requirement of considering the transverse aerodynamic force in the three-degree-of-freedom ballistic equation, the motion posture parameters of the deformed or locally damaged meteorites and debris fragments can be brought into a proper distribution statistical model for conversion;

step three, determining a spacecraft system/subsystem level disintegration criterion parameter and a condition boundary parameter; the method comprises the following steps of (1) providing a specific research object for analyzing and forecasting the meteorology process of the spacecraft in principle based on a three-level model of meteorology disintegration of the spacecraft; however, aerodynamic/thermal phenomena in the meteorology process of the spacecraft are actually an unsteady problem of continuous change of the appearance and the environment, so that the meteorology process is necessarily an uncertain process; in other words, assuming two merles with the same initial state, the merle process, survival condition of debris, and actual falling area will be different;

the analysis of the uncertainty process cannot be limited to the solution of the certainty problem, and therefore, a parameter statistical method based on condition boundaries is provided as a basic strategy for spacecraft meteor analysis and forecast; the parameter statistical method based on the condition boundary aims at flight environment, aerodynamic force/heat, flight motion and attitude parameters with uncertainty in simulation modeling analysis of the meteor process of the spacecraft, and determines the upper limit and the lower limit of the flight environment, namely the condition boundary according to the evaluation of the existing basic research results; constructing a proper parameter distribution mode in the boundary value domains of the conditions during analysis of the meteority process; constructing a disintegration condition criterion of structure disintegration damage; based on a statistical analysis correlation mathematical method, combining flight motion-force-heat-structure destruction disintegration calculation to carry out quantitative analysis on the concerned target parameters; obtaining statistically significant spacecraft flight path, disintegration process, debris viability and ground risk analysis evaluation results in the form of distribution bands with a certain degree of confidence;

from the aspect of flight motion, a three-degree-of-freedom ballistic equation can be adopted in general, but a six-degree-of-freedom ballistic equation can be adopted when the movement details before disassembly need to be mastered; considering the requirement of transverse aerodynamic force based on a three-degree-of-freedom ballistic equation, and taking the motion posture parameters of deformed or locally damaged meteorites and debris fragments into a proper distribution statistical model for consideration;

analyzing from the aspect of aerodynamic force, estimating one or more possible corresponding attitude angles of the static stable points according to the specific analysis of aerodynamic force characteristics of the moving attitude of the meteorole, and taking the static stable points as the basis of an attitude angle parameter distribution statistical model so as to carry out statistical quantification on the meteorole aerodynamic force; the three-degree-of-freedom ballistic calculation adopts a pneumatic force data table or a rapid engineering method as a main choice along ballistic calculation; the necessary trim attitude assessment is analyzed by numerical simulation; the fluctuation of the atmospheric parameters is a normal state; parameters needing to be included in the condition boundary parameter statistics comprise pneumatic coefficients and atmospheric density;

from the analysis of the aspect of aerodynamic heat, due to the uncertainty of the shapes of the deformed falling bodies and the debris fragments, no logical necessity exists for carrying out parameter distribution statistics on the directly related parameters of the aerodynamic heat; in practice the effect of aerodynamic heat can be translated into a consideration of the effect of aerodynamic uncertainty.

The analysis of the structure disintegration is the most difficult and the most lack of targeted basic research result support in the spacecraft merle analysis forecasting technology; the method is characterized in that a three-level model of the meteor disintegration of the spacecraft is combined, and analysis is carried out by establishing a disintegration condition criterion, namely relevant disintegration damage criteria are extracted according to results of a plurality of basic researches; the criteria for the disintegration conditions that may be considered include:

1) height criteria: that is, a certain type of disintegration process is set to occur at a certain height value; 2) melting criterion: namely, the metal material is set to have a certain form of disintegration process at the melting temperature; 3) and (3) pyrolysis criterion: namely, the composite material is set to be subjected to a certain form of disintegration process after being completely pyrolyzed at the pyrolysis temperature; 4) temperature criterion: namely, the object plane temperature is considered to be up to a certain set value and a certain form of disintegration process will occur; 5) heat flow criteria: that is, the accumulated heat flow or the instantaneous heat flow reaches a certain set value and a certain form of disintegration process occurs; 6) dynamic pressure criterion: namely, the dynamic pressure reaches a certain set value and a certain form of disintegration process is generated; 7) and (3) comprehensive criteria: the combination of the two or more criteria can include the weight sum, the achievement of any criterion, the achievement of all criteria and the like;

the melting criterion and the pyrolysis criterion can be regarded as a certain composite form of a temperature criterion and a heat flow criterion, and the melting criterion is listed to facilitate application analysis of the metal material with the largest proportion in the current spacecraft structural material; the pyrolysis criterion is used for carrying out application analysis on the main material composite materials of the containers with special requirements on the current spacecraft; when the above engineering criteria are adopted, the most important parameters related to structure disintegration and needing to be included in the condition boundary parameter statistics include: the size of the disintegrated fragments, the shape mode distribution of the disintegrated fragments and the range of the disintegrated height.

In the preliminary induction, the statistical analysis parameters related to the parameter statistical method based on the condition boundary for the analysis and forecast of the meteor of the spacecraft are as follows: (1) initial position (coordinate system XYZ or longitude/latitude/altitude) (2) initial velocity (three components of velocity or initial velocity/initial ballistic inclination/initial azimuth); (3) aerodynamic coefficient; (4) atmospheric density; (5) a disintegration height; (6) fragmentation property (material/shape/dimension); (7) and (5) breaking up the fragment gesture. The 7 sets of parameters can be further degenerated according to the situation and can be expanded if necessary. In order to use the 'parameter statistical method based on the condition boundary', the reference value, the uncertainty and the distribution form of the parameters need to be grasped, and the setting of the parameters related to the structure disintegration is most critical and difficult; the uncertainty of the parameters in the above (1) and (2), namely the condition boundary, is provided by a flight control and orbital transfer control system; (3) conditional boundary reference for group parameters the outcome of the aerodynamic study; (4-7) the uncertainty of the parameter is speculatively determined by the accumulation of experience and related observations;

specifically, the system/subsystem level disintegration criterion parameters and the conditional boundary parameters mainly include the above-mentioned group (1) to (5).

Fourthly, calculating and simulating the flight path of the spacecraft system/subsystem level; when the meteorology process of the spacecraft is analyzed, the grasp of the flight motion is a main clue running through the whole analysis process and is also a main content concerned by analysis, forecast and modeling; before the spacecraft is disassembled, namely, the system/subsystem level is aimed at, the accurate aerodynamic force/heat and motion analysis is carried out on the spacecraft, so that the flight parameters such as the height, the speed and the like at the time of the complete spacecraft breaking and disassembling are accurately deduced; if the mass characteristics of the system/subsystem level research object have very accurate information, a process of obtaining flight tracks and attitude changes can be solved in detail by adopting a six-degree-of-freedom ballistic equation;

fifthly, decomposing, destroying, calculating, simulating and analyzing the hierarchical structure of the spacecraft system/subsystem; wherein, the analysis result of aerodynamic force and aerodynamic heat is used as the boundary condition of the object plane of the structural stress analysis; obtaining physical property parameters changing along with temperature and thermal stress caused by temperature difference by thermal analysis in the structure; under the comprehensive action of object surface load transfer stress and thermal stress, analyzing structural damage according to the total stress distribution condition of the material and parameters such as Young modulus, Poisson ratio and the like under corresponding conditions;

the mass loss rate of the metal material is determined by an energy balance equation by adopting a melting point control calculation model; the heat transfer calculation of the metal material is solved by a general solid heat transfer equation, and the radiation heat transfer is usually considered besides the convection heat transfer in the boundary condition; the melting point control calculation model is a relatively conservative model, and the melted material can be judged as a reserve in theory; however, our research results show that the time for raising the surface temperature of the aluminum alloy cabin from the softening temperature to the melting temperature is in the order of milliseconds, and the time for completely melting the aluminum alloy cabin with the size of one meter is in the order of tens of milliseconds. In summary, when the spacecraft falls into the atmosphere, the melting point of the aluminum alloy material which is often used as a main structural component is low, and an analysis conclusion applicable to engineering can be obtained by adopting a melting point control model, so that the spacecraft is completely acceptable in engineering. For high-melting-point alloys adopted by special parts, such as titanium alloy, niobium alloy, tungsten alloy and the like, the temperature rise and melting conditions of the high-melting-point alloys are mainly considered, and whether the debris exists or not is judged;

the reaction of carbon-based composite materials commonly used in aerospace engineering at high temperature mainly takes carbon-oxygen combustion, carbon-nitrogen combination and solid carbon sublimation into consideration; in the calculation of the heat transfer of the carbon-based material, heat source items such as heat clearing heat, gasification heat and the like are considered on the basis of a heat conduction equation; under given pressure and temperature, iteratively solving component concentration and mass loss rate according to a chemical reaction formula and compatibility conditions, and solving the material ablation rate after the mass loss rate is solved; when the ablation rate of the material is calculated, the pressure is given by the outer edge parameters of the boundary layer, the surface temperature is determined by a surface energy balance equation, and the boundary condition of the boundary layer and the internal heat conduction is determined by the surface energy balance equation;

sixthly, determining a spacecraft component level disintegration criterion parameter and a condition boundary parameter; wherein, the determination of the component level disintegration criterion parameter and the condition boundary parameter refers to the specific method of the system/subsystem level in the step three;

specifically, the parameters concerned by the component-level disassembly criterion parameters and the conditional boundary parameters include: the component level initial position (coordinate system XYZ or longitude/latitude/altitude), the initial velocity (three components of velocity or initial velocity/initial ballistic inclination/initial azimuth) inherit the parameters at the moment of system/subsystem level disassembly; the rest of the components comprise aerodynamic coefficient, atmospheric density and disintegration height;

seventhly, calculating and simulating a flight path of the spacecraft component level; because the shapes of the parts or fragments of the spacecrafts after meteorites are generally complex, the quality characteristics and the shapes are still in continuous change, and the flight motion attitude and the related aerodynamic force/thermal action have great uncertainty. For the problems, obtaining the attitude change process of the meteorolite part or fragment is difficult and has no sufficient engineering significance, so that a rapid calculation method combining a three-degree-of-freedom ballistic equation with aerodynamic heat is adopted for calculating and simulating the flight path of the part level; calculating attenuation change of the mass of a research object according to the fusion ablation rate in the simulation process;

eighthly, calculating, simulating and analyzing the decomposition and damage of the hierarchical structure of the spacecraft component; the method of component hierarchy structural disintegration damage computational simulation analysis is substantially the same as the system/subsystem hierarchy; but a more rapid simplified model can be adopted in the aerodynamic and aerodynamic heat treatment links;

constructing a spacecraft fragment/particle level block strip model; in the shape modeling of the fragment/particle level, the geometric characteristics of the fragment/particle level are highly abstracted; the geometric modeling abstraction is based on the principle of a complete induction method, and different geometric characteristics can be covered, so that corresponding pneumatic characteristics and flight motion characteristics are covered and embodied; the specific method is that the fragments/particles are divided into three types according to geometric characteristics: block (Block); ② (Column); slice (Slice); therefore, this division mode of the Slice/grain level is called a blockslice Model (BCSM: Block-Column-Slice Model);

in the geometries of the three types of chips/particles, the "block" is designated as a sphere or ellipsoid and is shaped by the minor-major axis ratio; "sheet" is referred to as a semi-circular edge disc, shaped by the thickness to diameter ratio; the strip is indicated by a cylinder with a hemispherical end face and is shaped by a slenderness ratio; the scale ranges (upper and lower limits) of the three types of fragments/particles are set according to requirements, and the block is determined by a characteristic scale long axis; the "bar" is determined by the characteristic dimension column length; "patch" is defined by a characteristic dimension diameter; the geometrical shape of the debris/particles is set to be round and smooth in surface, and the debris/particles generally do not have the characteristic of distinct edges and corners after being subjected to airflow scouring based on meteorite debris;

the chips/particles need to meet material (density) and corresponding total mass constraints; the proportion of the fragments to the total mass and the proportion of the fragments/pieces/strips and the specific setting of the related parameters of the three types of fragments/particles are presumed according to past experience and observation results, and meanwhile, the method also depends on the cognition degree of basic research results on the meteorite disintegration process;

in some cases, if the system/subsystem level integer generator undergoes a violent disassembly process (such as container or battery burst), the component level may not exist, and the component level directly evolves from the system/subsystem level to the fragment/particle level; in the presence of the component level, some debris/particulate entities may also be directly generated by the disintegration of the monolith (mainly the shell);

table 1 defines the chip/particle geometry and material property parameters.

TABLE 1 chip/particle geometry and Material Property parameters

For each particular chip/particle, the individual number of geometric and material property determinations can be expressed as:

n＝n(T,R,L,M)

in the above parameters, T and M are discontinuous variables, R and L are continuous variables, and in particular, the number of individuals corresponding to the block, bar, slice type can be respectively expressed as:

n_B＝n(B,R,L,M)；n_C＝n(C,R,L,M)；n_S＝n(S,R,L,M)

the continuity variables R and L may be generally expressed in the form of distribution functions within a certain range; discretization treatment is carried out as required in practical application; based on the trend that the molten ablation and airflow scouring action in the meteorite process have the effect of enabling the shape of the debris to be round, the lower limit of the value of the setting parameter R is recommended to be 1, and the upper limit is evaluated and determined according to the basic research and the experience accumulation of the observation result; in addition, when the setting parameters of the three fragment types are 1, the three fragment types are actually classified into reference spherical shapes; the characteristic dimension L must have definite upper and lower limits, which are determined according to the concerned problem and the condition evaluation of the actual research object; any type of fragments with the same material property should meet the constraint condition that the sum of the masses of the same material is unchanged;

the key parameter of the fragment quantity is the characteristic scale (or fragment quality), and similar researches show that the fragment quantity and the characteristic scale are similar to an exponential change rule, namely, the fragments with smaller scale have larger quantity. For this purpose, an expression of the number of fragments with a characteristic scale larger than L is constructed as follows:

in the formula N₀Total number of fragments to be included in the statistics, b is the disintegration intensity coefficient, L₁And L₂Respectively the lower limit and the upper limit of the characteristic scale of the fragments, and the value of L is between L₁And L₂To (c) to (d);

the 'block strip model' of the chip/particle level can be subjected to necessary degeneracy in terms of category, shape, characteristic dimension, material and the like according to the engineering requirements of the actual problem.

Tenthly, analyzing the survivability of the spacecraft fragments and the boundary of a simulation calculation condition; before performing the debris fraction viability analysis, a proper definition or definition of the so-called debris fraction is required; the term "debris" refers to a part of a part from which a part belonging to the part can be clearly seen; whereas fragments (or particles) refer to residues whose original morphology is not clearly visible. Survival means falling to the ground;

according to calculation, the limit equilibrium speed of solid aluminum particles with the diameter of 1mm falling to the earth surface is about 4m/s, so that the millimeter-scale fragment particles have obvious pain or injury when impacting a human body; this can be used as a reference value for the lower limit of the geometric scale when we analyze fragment viability. The standard of gun identification, which is specified by the ministry of public security of China, is 1.8 joules/square centimeter, and can be used as another reference value of the lower limit of the kinetic energy of the fragments;

quantitative analysis of debris viability assessment relies on empirical observations and the accumulation of outcomes from basic research;

as a product of the disintegration, the parameters considered by the fragment simulation computation condition boundary analysis include all the parameters listed in step three, wherein the fragment-level initial position (coordinate system XYZ or longitude/latitude/altitude), the initial velocity (three components of velocity or initial velocity/initial ballistic inclination/initial azimuth) inherit the parameters of the component or system/subsystem-level disintegration time;

step eleven, calculating and simulating the aerodynamic force and flight path of the spacecraft fragment/particle level; the research object of the fragment/particle level is highly abstracted and modeled, namely, the aerodynamic force and the flight path of the research object are calculated and simulated according to the block strip model; the aerodynamic force adopts aerodynamic force coefficient tables of different shapes for carrying out interpolation; and calculating and simulating the flight path by adopting a three-degree-of-freedom trajectory equation. In the calculation simulation of fragment/particle level aerodynamic force and flight path, the method provided by a parameter statistical method based on condition boundary is mainly considered in an important way to obtain quantitative landing area distribution data;

twelfth, analyzing and evaluating a falling area of the spacecraft fragment/particle level; and analyzing and evaluating the landing area dispersion according to the flight path data of the debris fragments obtained in the step eleven. The landing zone range depends on the following factors: 1) flight parameters and uncertainty at the moment of disassembly of the whole device and the components; 2) the material, shape, size and distribution rule of the debris; the statistics of the landing area distribution can adopt a traditional Monte Carlo point method and can also adopt a more efficient modern experimental design algorithm.

Thirteen, ground risk assessment; evaluating the ground risk degree by adopting a casualty area and total casualty probability calculation method provided by a literature according to the calculated results of the fragment survivability, the flight track and the landing area obtained in the tenth to twelfth steps; it is defined as follows:

step 13.1 area of injury and death

Equivalent area of injury and death A_cFor assessing ground risks posed by single reentry event debris, it is a combination of debris cross-sections and individual projected cross-sections. The total area of injury or death is determined by the sum of all n pieces in a reentry event:

in the formula A_hIs the projected cross-sectional area of the human ground, A_iIs the maximum cross-sectional area of a single piece of debris falling to the ground. Human body ground projection cross section A used in the NASA Security Standard NSS 1740.14_h0.36 square meters (m2) (mean values from human body shape statistics). According to the NASA safety standard, total casualty area of less than 8 square meters (m2) is safe.

Step 13.2. Total casualty probability

The overall casualty probability E represents the number of people who may be hit by reentry debris:

ρ_iis the population density at the ith fragment drop point, which can be estimated on average according to urban and rural locations, in detail as the distribution data of the earth population density (2005 year) published by Columbia university;

is the area of injury or death of the ith fragment; sigma_iIs the probability of the ith fragment hitting a unit area, calculated by:

delta is the latitude of the location of the landing point,. phi is the angle of inclination of the re-entry orbit, R_EIs the radius of the equivalent sphere of the earth. When the total casualty probability E is less than 1:10000, the risk of debris reentry is considered acceptable.

Example 1:

the method for analyzing and forecasting reentry and meteority of the spacecraft comprises the steps shown in figure 1, and is explained through an embodiment 1 in order to more clearly explain the technical scheme of the invention. The embodiment 1 is used for analyzing and forecasting the situation similar to the situation that the China freight ship enters the air space after the mission is over, and because the related basic research supporting work is a long-term continuous process, the boundary and distribution setting of some parameters in the split model are continuously refined and enriched through the basic research result database.

Analyzing a spacecraft structure and constructing a disintegration three-level model; FIG. 2 is the spacecraft for reentry merle analysis, which is a two-compartment structure, the two compartments being a sealed compartment and a resource compartment, respectively; the cabin body is made of aluminum alloy and is tightly connected by steel bolt flanges, so that the two cabin assemblies can be considered as a whole; solar panels (solar sailboards) are arranged on two sides of the resource cabin, and a relay antenna assembly is arranged on the upper part of the resource cabin; in addition, other assemblies expose out of the cabin body, including rail control and attitude control engine spray pipes, various optical device lens windows, and the like; the two-cabin assembly formed by the complete spacecraft and the falling-off solar sailboard of the spacecraft belongs to the system/subsystem level of the three-level model of the meteorology disintegration of the spacecraft in the embodiment analysis;

the spacecraft reenters the component level of the merle disintegration three-level model, which roughly comprises the ten types in FIG. 3; they are the core functional components that make up the spacecraft; most of the internal components are made of aluminum alloy materials, and a small amount of stainless steel components; the rail control and attitude control engine spray pipes are made of high-temperature-resistant alloy; also a bottle and can type container partially made of composite material or carbon fiber; in addition, a battery module is also provided;

the spacecraft enters the debris/particle level of the meteorite disintegration three-level model again, the simplification is set as spheres, cuboids, sheets, rods (slenderness ratio is 10) and rods (slenderness ratio is 5), and statistical calculation analysis is carried out by adopting a grouping method on the basis;

for systematic modeling and simulation of merle disintegration process, the following needs to be grasped: determining meteor initial conditions; the timing and symmetry of the shedding of the solar panels; the complexity of gradual melting and disintegration of the two-compartment assembly; uncertainty in local parameters of deformation and aerothermal during component survival; the possibility and effect of physical or chemical explosion of the fuel tanks and the cell modules; the uncertainty of the above situation makes the condition parameter setting principle of meteor analysis for disassembling the whole device adopt conservative setting as much as possible, including comprehensive evaluation and analysis of the condition state parameter reference value and the value range thereof in the possible range of the engineering object;

secondly, calculating and analyzing system/subsystem, component level aerodynamic force, aerodynamic heat and flight characteristics; according to the known reentry parameters and possible flight parameter ranges, comprehensive calculation and analysis are carried out on the aerodynamic force, the aerodynamic heat and the flight characteristics so as to obtain the basic aerodynamic characteristics and the flight characteristics of the meteor spacecraft system/subsystem and component level. The computational analysis of this step is based on preset possible flight conditions;

step 2.1, aerodynamic force evaluation; the system/subsystem level has a determined shape before the disintegration occurs, but mass aerodynamic performance data under different heights, different speeds, different flow states and different postures are considered in the whole flight process, and only a rapid engineering calculation method is used. Therefore, the system/subsystem level aerodynamic force evaluation is mainly based on an engineering method, and for some typical states, results of a numerical simulation method are compared, verified or corrected;

the aerodynamic engineering calculation method adopts a surface element method, and comprises Newton theory and correction of continuous flow region, molecular collision theory of free molecular flow region, and overlapping formula processing of thin transition flow region;

the first step of the hypersonic surface element method is preprocessing the geometric shape of the aircraft, dividing an object plane into a plurality of micro surface elements, approaching a complex shape by utilizing a quadrangle or a triangle, and dividing and calculating simpler objects such as spherical, conical and cylindrical geometric shapes by using an analytical method; for complex shapes, mesh segmentation is generally performed by means of professional software; the aerodynamic engineering calculation requires the calculation of the coordinates of the central point of each surface element, the normal vector of the surface element representation and the surface element infinitesimal area;

calculating the pressure coefficient of an object surface element by a surface element method, solving the moment of the object surface element in a coordinate system, and summing to obtain the aerodynamic force of the aircraft;

the aerodynamic performance numerical simulation method for verification, comparison and analysis comprises Boltzmann model equation unified algorithm, DSMC numerical simulation, N-S/DSMC coupling algorithm, NS equation numerical simulation and the like, and can also perform some wind tunnel force measurement test analysis to confirm the aerodynamic performance algorithm when necessary;

FIG. 4 is a diagram of a typical six-component parameter of aerodynamic performance as a function of angle of attack for a system/subsystem level spacecraft integer during reentry flight; corresponding to the performance curve under the condition of a specific height attitude; the parameter state combination required in the actual engineering problem analysis comprises data packets under various condition states; FIG. 5 is some results of numerical simulations;

step 2.2, evaluating pneumatic heat; in the embodiment, the aerodynamic heat evaluation of the reentry and meteorology process of the large-scale spacecraft mainly obtains thermal environment parameters under the condition of reentry flight parameters, and provides basis for the subsequent structural disintegration analysis; for the thermal environment computational analysis of the system/subsystem hierarchical configuration, the main work carried out includes: (1) performing aerodynamic heat calculation on a spacecraft complex-shape shaper and an irregular disassembly part; (2) predicting the heat flows of the local protrusions and the cavities on the surface, performing pneumatic heating calculation of the local protrusions and the cavities on the surface, and calculating the heat flows and peak heat flows of an interference area and a reattachment area through an engineering fitting formula;

aiming at the integrated calculation requirements of aerodynamic force, heat, ablation and trajectory of meteorology forecast of a large spacecraft, a method for calculating the aerodynamic heat of a whole spacecraft and a disassembly part based on a non-structural grid is adopted, and the coupling calculation with a pneumatic ablation module is realized through a calculation interface coupled with the pneumatic ablation;

under the unstructured grid, a ball head stagnation point is calculated by adopting an Fay-Riddell formula, a three-dimensional effect is considered for the three-dimensional stagnation point, and dimensionless heat flow in a non-stagnation point area is approximated by a sine function of an impact angle; for free molecular flow, a heat flow calculation formula given by a molecular motion theory is adopted, and a bridge relation method is adopted in a thin transition flow region;

adopting a method for calculating the stagnation point parameters including the stagnation point position, the stagnation point type and the stagnation point radius of the nonspecific shape under different attack angles; namely, for the same component, the position and the type of the stagnation point are different under different incoming flow states, for example, when the incoming flow attack angle of the gas cylinder component is 0 degree, the stagnation point is a ball head; when the incoming flow attack angle is 90 degrees, the stagnation point is a cylinder;

the calculation of the pneumatic thermal environment of the bulge adopts two technical schemes: for the bulge smoothly connected with the body part, the integral calculation of the bulge and the body part is realized by a spline fitting method; for the convex objects with unsmooth and continuous shapes, calculating the heat flow of the convex objects, the peak heat flow of the interference area and the reattachment area by utilizing an engineering fitting formula; for the calculation of the aerodynamic thermal environment of the concave cavity, the concave cavity and the gap are mainly used as independent parts for calculation, the calculation is divided into a closed concave cavity and an open concave cavity by combining the geometric appearance, the flow characteristics and the local flow field parameters, and the peak heat flow of a separation interference area and a reattachment area is calculated;

FIG. 6 shows the result of an aerodynamic heating calculation at a certain condition state at the system/subsystem level; the system level subject here, the re-mersible spacecraft integer, is in a configuration before any disassembly has occurred; the subsystem-level research object is the configuration of two cabin bodies after the solar sailboards on the two sides of the resource cabin fall off; FIG. 7 shows the results of the aerodynamic thermal calculations for a portion of the part level study; the parts comprise an orbit control engine spray pipe, a relay antenna and a main bearing module; the actual analytical assessment work was calculated for all the parts contained in merle (see FIG. 3);

step 2.3, evaluating the flight motion; the flight motion evaluation of the system/subsystem level and component level research objects is mainly simulated and solved by a three-degree-of-freedom ballistic equation;

the reentry of the spacecraft has a determined shape before the disintegration occurs, and theoretically, a six-degree-of-freedom ballistic equation can be used for solving flight parameters of the spacecraft, namely, the change of position parameters and attitude parameters of the spacecraft along with time is obtained, and then each component of the flight speed and each parameter under a relevant coordinate system relative to an observer or the earth, such as longitude and latitude, height, ballistic inclination angle and the like, are derived. However, for the following reasons: 1) the system/subsystem level aircraft is positioned in a free molecular flow or rarefied transition flow region, and the attitude and the like in the reentry process have weak influence on the position and speed key parameters; 2) the precise mass characteristics of such aircraft in reentry flight, including the center of mass and moment of inertia, are difficult to ascertain; therefore, the calculation result of the three-degree-of-freedom ballistic equation used in engineering can also provide applicable data; the biggest problem faced by the component level is incomplete certainty of the real shape of the component, namely although the predefined component shape is relatively determined, the actually decomposed component configuration is difficult to avoid connecting structure pipelines or tearing, breakage and the like, so the flight evaluation of the component level also mainly adopts a three-degree-of-freedom ballistic equation to carry out simulation solution;

FIG. 8 is the system/subsystem flight motion calculation results under two different initial conditions, from which the height and velocity parameters of possible meteor disintegration can be preliminarily evaluated, laying the foundation for subsequent disintegration analysis;

step three, determining a system/subsystem level disintegration criterion parameter and a condition boundary parameter; in the analysis of melting, pyrolysis/ablation disintegration of structural materials when the spacecraft with the complex structure enters the atmosphere again, the disintegration criterion is the key of damage analysis; for the metal material, a melting point temperature control model is adopted, namely, the surface of the metal material is supposed to reach a melting point, and the surface of the material is melted and lost; for the carbon-based composite material, a pyrolysis/ablation control model is adopted, namely the structure is supposed to be completely pyrolyzed or ablated to be damaged; the system/subsystem hierarchy relates to conditional boundary parameters including: (1) initial position (coordinate system XYZ or longitude/latitude/altitude) (2) initial velocity (three components of velocity or initial velocity/initial ballistic inclination/initial azimuth); (3) aerodynamic coefficient; (4) atmospheric density; in the embodiment, the range of a simulation result of basic coverage can be obtained by taking the values of parameters under two extreme attitude conditions of 0-degree attack angle and 90-degree attack angle of a spacecraft whole device (with a solar sailboard) and two capsule bodies (without the solar sailboard) in the rapid analysis;

fourthly, calculating and simulating a system/subsystem level flight path; on the basis of the flight motion evaluation in the step 2.3, the condition boundary parameters set in the step three are combined, aiming at one condition of spacecraft reentry parameters (longitude, latitude, altitude, speed, trajectory inclination angle and azimuth angle) in the embodiment, the flight tracks of the whole spacecraft at the system level and the two cabins at the subsystem level are respectively calculated and simulated, and the calculation result is shown in a figure 8;

fifthly, decomposing, destroying, calculating, simulating and analyzing the hierarchical structure of the system/subsystem; in combination with the outline and structure of the system/subsystem level aircraft shown in fig. 2, the solar panel material is mainly a composite material, and when the temperature of the composite material reaches above 150 ℃, the base material of the composite material undergoes vitrification softening and conversion, and the mechanical property is reduced. The two cabin bodies are mainly made of aluminum alloy (density 2800 kg/m)³Specific heat 921J/(kg. K), thermal conductivity 121W/(m. K), melting temperature 933 (K)); according to "step 2.2. aerodynamic heatingEvaluating the obtained surface thermal flow data at different heights along the flight trajectory, and performing structural heat transfer, fusion or ablation calculation on the system/subsystem level; FIG. 9 shows a cloud of surface temperature profiles under some exemplary conditions; the temperature of the windward front edge of the solar panel frame is highest, and the position of the butt joint is the position of the butt joint; the amount of ablation of the windsurfing board material is small, but according to the temperature that its bonding material can withstand, it is judged from the pyrolysis performance that it will undergo a first disintegration (frame breakage of the links between the windsurfing boards) during a height of 110km to 105 km; after the solar sailboard is broken, the posture of the two cabins is influenced by the new aerodynamic performance, and the two cabins tend to be re-balanced to fly at an attack angle of 170 degrees (the resource cabin is in front); in order to obtain accurate temperature information of the two cabin bodies, heating before posture adjustment needs to be considered when the surface temperature of the two cabin bodies is calculated; the surface temperature distribution of the two chambers in FIG. 9 at a height of 100km indicates that the maximum back-off for the bottom ablation has exceeded 1mm, and it is judged that it will disintegrate at this height; when the height of the relay antenna is 100km, the temperature of the structural material of the relay antenna does not reach the melting point, and the relay antenna is decomposed along with the two cabins and then is used as a component level to continue falling into the atmosphere; in order to further confirm the melting and disintegration conditions of two cabin bodies (large-size and thin-shell structures) of the spacecraft with the complex structure under the high-temperature condition, a three-dimensional finite element heat transfer model is applied, coupling solution is carried out on aerodynamic heat and structural thermal response, and high-precision three-dimensional numerical calculation is carried out on the temperature field under the transient aerodynamic heat environment and material melting condition at certain typical heights when the two cabin bodies reenter the flight along the trajectory, so as to verify and supplement the calculation result of the rapid engineering;

sixthly, determining component level disintegration criterion parameters, conditional boundary parameters and flight path calculation simulation; the disintegration criterion of component level merle is similar to the system/subsystem level; for the metal material, a melting point temperature control model is adopted, and the surface of the metal material is supposed to reach a melting point, so that the surface of the material is melted and lost; for the carbon-based composite material, a pyrolysis/ablation control model is adopted, and the structure is supposed to be completely pyrolyzed or ablated and damaged; the component hierarchical structure material mainly comprises aluminum alloy, and also comprises partial stainless steel, carbon fiber composite material and a small amount of high-temperature-resistant alloy material; the component level comprises some containers and accumulator modules, which may burst during the disassembly process; the conditional boundary parameters involved at the component level include: (1) initial position (coordinate system XYZ or longitude/latitude/altitude) (2) initial velocity (three components of velocity or initial velocity/initial ballistic inclination/initial azimuth); (3) aerodynamic coefficient; (4) atmospheric density; (5) system/subsystem level of disintegration height; wherein the component level initial position and initial velocity inherit the parameters at the time of system/subsystem level disassembly; similarly, in the embodiment rapid analysis, the minimum windward area (or attack angle 0 ℃) and the maximum windward area (or attack angle 90 ℃) of the component are taken as values under two extreme conditions, so that a simulation result range of basic coverage can be obtained; FIG. 10 shows the results of the calculation of flight paths for some components, including the system/subsystem level results and their relationship to the component level results;

step eight, decomposing, destroying, calculating, simulating and analyzing the component hierarchical structure; starting from the disassembly of the whole device (system level) or two cabins (subsystem level), the independent flight process of the structural functional components connected inside or outside the cabin is started; the component hierarchy inherits the relevant flight parameters during disassembly, and simultaneously considers the value range of the conditional boundary parameters given in the previous step during statistical calculation and analysis, namely the uncertainty of the parameters; taking the rail-controlled engine nozzle part as an example, the surface temperature distribution of the rail-controlled engine nozzle part at different heights is shown in FIG. 11. A melting point control model is adopted, namely, the metal is melted and absorbs heat after the melting point is reached, the surface temperature is kept at the melting point, and the melting and disintegration conditions of the part can be calculated, analyzed and evaluated; fig. 11 also shows the theoretical melting back distribution of the relay antenna at a height of 70 km; FIG. 12 shows the name, material and disintegration of each study object at the system/subsystem level and component level during reentry and merry of the spacecraft; the calculation and analysis of the embodiment show that when the large-scale spacecraft flies below 120km, the aerodynamic force/heat has obvious effect; at a position above 105km, the solar sailboard falls off and disintegrates under the action of aerodynamic heat; in the range of about 100-95 km in height, the shell of the two-cabin assembly body made of aluminum alloy will melt and disintegrate, and the internal parts begin to bear strong aerodynamic force/heat; respectively melting and decomposing the aluminum alloy and the stainless steel parts in a range of about 95-70 km in height; the carbon fiber and composite material parts are pyrolyzed sufficiently and disintegrated. High-temperature resistant alloy components such as a rail control engine unit and an attitude control engine unit can be completely melted at the height of about 60km, and the possibility that the remains of identifiable original forms fall to the ground is extremely low; in addition, the two cabins are disintegrated, and the battery module is rapidly cracked and disintegrated due to temperature rise under the action of pneumatic heat;

step nine to step twelve, analyzing a fragment/particle level model, a flight path and a falling area range; the fragment/particle level adopts a 'block and strip integrated parameterized model' to carry out geometric construction and statistical analysis; once the existing debris/particles are defined, they are considered to be reducible to the ground, no longer taking into account the aerodynamic thermal effects; scattering the falling region of the debris, and evaluating according to the related parameters of the actually fallen dead points; the landing zone range depends on the following factors: 1) flight parameters at the moment of disassembly of the whole device and the components; 2) the material, shape and size of the debris; in the embodiment, the debris/particle level of the meteorite three-level model is simplified and set to be a sphere, a cuboid, a sheet, a rod (slenderness ratio is 10) and a rod (slenderness ratio is 5), and statistical calculation analysis is carried out on the basis by adopting a grouping method; FIG. 13 is a calculation result of a range of falling points where debris may be generated under certain set conditions; based on the foregoing re-entry merle analysis, the summary of the relevant conclusions is as follows: the spacecraft is subjected to a plurality of complex atmospheric environments such as free molecular flow, rarefied transition flow, slip flow, continuous flow and the like in sequence in the falling process, and is subjected to complex aerodynamic force/thermal influence, the mechanical property of a structural material is gradually reduced, a metal truss structure deforms, softens and fails, and a composite material is pyrolyzed and ablated until the composite material is disintegrated; it is basically burnt out in the process of reentry of meteorite; if unburned chips are present, their sources may be niobium hafnium alloy orbital jet tubes and carbon fiber wound metal liner cylinders, followed by three larger scale optical assemblies. Thus, the conclusion on the survivability of the debris fragments in terms of probability statistics, the primary tone is "complete burn", the case of landing of debris fragments is a small probability event;

thirteen, ground risk assessment; performing calculation analysis on the basis of the certainty of the existence of the debris in the probability as a premise in the following ground risk assessment; according to the flight orbit characteristic of the relative relation between the spacecraft and the earth, the falling area range of the near merle period can be approximately predicted; the orbit subsatellite point can not cover the whole area between the south latitude and the north latitude of 42 degrees within a certain period, but basically passes through partial areas periodically; FIG. 14 is the position of the sub-star point in the remaining 138 revolutions; the tracks can be seen to be in a net structure; the grid void area in the figure, i.e. the safety zone during this period (8 to 9 days), considering the small lateral extent of the debris landing zone; as the distance from the meteor decreases, the number of remaining circles decreases, and the coverage of the sub-stellar dot decreases. FIG. 15 is the last 5-turn trajectory based on prediction; it can be seen that, at least half of the area within the range of 42 degrees of north and south latitude is in the safe range; according to the operational parameter characteristics of the orbit subsatellite point sweeping earth surface, when the orbit inclination angle is 43.8 degrees, the probability that a unit area at the earth surface is hit by a remaining debris/fragment is shown in fig. 16; according to the calculated results of the fragment survivability, the flight path and the landing area obtained in the tenth step to the twelfth step, a casualty area and total casualty probability calculation method provided by a literature is adopted;

considering only the components remaining, taking a rail-controlled engine as an example, assuming that the inclination angle of the trajectory is about 90 degrees when the rail-controlled engine falls to the ground and the maximum cross-sectional area is about 0.2 square meter, the total casualty area

So A_c＜8m²The whole is safe; if both parts and fragments are considered, even if the fragment area is negligible, when the number of fragments considered exceeds 22, there will be a according to the casualty area formula_c＞8m²I.e. not sufficiently safe; this represents a comprehensive quantitative effect of the number of fragments on the total area of casualty; the more accurate calculation of the total casualty probability requires integrating the drop point area of each fragment, so that in principle, the area division is required according to the global population density distribution dataPerforming score evaluation; due to the influence of the track inclination angle and the earth shape, the probability of hitting the unit area at different latitudes is different; however, in the preliminary estimation, the probabilities can be considered consistent, namely, the debris is completely randomly distributed between the north and south latitude 42 degrees; in the area, the land area accounts for about 28%, the population dense area (40-500 persons/square kilometer) accounts for about 10%, and the average value is 100 persons/square kilometer; particularly dense areas of the population (>500 persons/square kilometer) accounts for about 1%, and the average value is 1000 persons/square kilometer; as a recheck of the human mouth density value, the area of the earth between the south latitude and the north latitude 42 degrees is 3.4 multiplied by 108 (km)²) Then the population in the area is 6.8 × 109; the actual total population of the earth is 72.08 hundred million (2014), namely the estimated population of the region accounts for about 95 percent of the global population, so that the method is basically reasonable and conservative; maximum cross-sectional area of the chips as A_i＝0.01m²Meter, then

If the person falls into a dense population area, the population density is 100 persons/square kilometer, and the casualty probability caused by a single fragment is as follows:

if the population falls into a particularly dense region, the population density is 1000 people/square kilometer. The probability of casualty caused by a single fragment is:

fall into other areas, considered safe, disregarded; the combined casualty probability of a single fragment is: p_i＝10％·P_i1+1％·P_i2＝9.8×10^-6(ii) a Since the probability of casualty of a single fragment is small, in the case where the total number of fragments is not so large, the total probability of casualty caused by a plurality of fragments is: 1- (1-P)_i)^N≈N·P_i(ii) a According to the NASA safety standard, the risk is acceptable when the total casualty probability is below 1: 10000. I.e. P ≈ N.P_i＜1×10^-4N is less than 10^-4/9.8×10^-610.2; namely, the number of fragments with the diameter larger than 0.1m is required to be not more than 10 when the fragments fall to the ground; if it isConsidering only the rail-controlled engine parts, the maximum cross-sectional area is about 0.2m²Total area of injury and death 1.1m²The total casualty probability (4 parts) is about 1: 11400; the term total casualty probability is used herein to refer to the mathematical expectation of the total number of casualties caused by debris/debris. For example, the probability of casualty of a single fragment is 9.8 × 10^-6(≈ 1:102000) indicating: one fragment can cause one-hundred-thousand casualties on the earth, or the probability of one-hundred-thousand casualties in the world is one-hundred-thousand, or roughly speaking, one-hundred-thousand fragments can cause one-hundred-thousand casualties; in particular, the probability that a particular person is hit by a particular fragment should be divided by the total number of people worldwide, i.e., about seven billion; by contrast, the probability of a two-color jackpot for a particular person is 1:1772 ten thousand; the probability of the middle and large lotto jackpot is 1:2142 ten thousand; the probability of a total damage accident caused by one flight when the aircraft is collided is about 1:530 ten thousand;

while embodiments of the invention have been described above, it is not limited to the applications set forth in the description and the embodiments, which are fully applicable in various fields of endeavor to which the invention pertains, and further modifications may readily be made by those skilled in the art, it being understood that the invention is not limited to the details shown and described herein without departing from the general concept defined by the appended claims and their equivalents.

Claims (9)

1. An analysis forecasting method for reentry of spacecraft into meteor space is characterized by comprising the following steps:

analyzing a spacecraft structure and constructing a disintegration three-level model; the disassembly three-level model comprises a spacecraft system/subsystem level, a spacecraft component level, and a spacecraft debris/particle level;

calculating and analyzing aerodynamic force, aerodynamic heat and flight characteristics of a spacecraft system/subsystem level and a spacecraft component level;

step three, determining a spacecraft system/subsystem level disintegration criterion parameter and a condition boundary parameter;

fourthly, calculating and simulating the flight path of the spacecraft system/subsystem level;

fifthly, decomposing, destroying, calculating, simulating and analyzing the hierarchical structure of the spacecraft system/subsystem;

sixthly, determining a spacecraft component level disintegration criterion parameter and a condition boundary parameter;

seventhly, calculating and simulating a flight path of the spacecraft component level;

eighthly, calculating, simulating and analyzing the decomposition and damage of the hierarchical structure of the spacecraft component;

constructing a spacecraft fragment/particle level block strip model;

tenthly, analyzing the survivability of the spacecraft fragments and the boundary of a simulation calculation condition;

step eleven, calculating and simulating the aerodynamic force and flight path of the spacecraft fragment/particle level;

twelfth, analyzing and evaluating a falling area of the spacecraft fragment/particle level;

thirteen, ground risk assessment;

the spacecraft system in the spacecraft system/subsystem level refers to a whole spacecraft form, and the spacecraft subsystem is a relatively complete group section of an oversize spacecraft or a complete cabin section with external accessory parts falling off; the spacecraft component level refers to an external object or an internal object which has certain functionality and structure shape certainty inside and outside a spacecraft cabin; the spacecraft debris/particle level refers to the residue of the original geometric form which cannot be visually identified, and belongs to an abstract model.

2. The method for analyzing and forecasting reentry of spacecraft into meteor space of claim 1, wherein in the second step, the aerodynamic force calculation analysis adopts an engineering calculation method or a numerical simulation method; based on the characteristics of the meteorolite and the complex shape of the wreckage of the spacecraft, the aerodynamic force calculation is mainly based on an engineering method; performing supplement and comparison and verification of typical state calculation results by using numerical simulation;

in the second step, the pneumatic thermal parameters before the whole system/subsystem level spacecraft is disassembled are obtained by numerical simulation such as a DSMC method; from the engineering application angle of spacecraft meteor analysis and forecast, a rapid engineering calculation method is adopted along a large number of thermal environment parameters of a trajectory; due to the inexact expectation of the shape of the deformed falling body and the disassembled part, the local details are properly simplified in the calculation of the thermal environment engineering.

3. The method for analyzing and forecasting reentry of spacecraft into meteor space as claimed in claim 2, wherein the aerodynamic engineering calculation method adopts a binning method, including continuous flow region Newton theory and its correction, free molecular flow region molecular collision theory, and rarefied transition region lap joint formula processing; the numerical simulation method comprises any one of Boltzmann model equation unified algorithm, DSMC numerical simulation, N-S/DSMC coupling algorithm and NS equation numerical simulation;

in the specific processing of the pneumatic thermal parameter acquisition, the meteor windward side is subjected to axisymmetric matching assumption, and the problem of streaming with an attack angle and a sideslip angle is converted into the problem of simple total attack angle streaming by adopting an equivalent method; the problem of zero-attack-angle streaming can be further solved by applying an equivalent spherical cone method; after the transformation, the aerodynamic heat problems of an attack angle and a belt sideslip can be calculated by using a zero-attack-angle spherical cone aerodynamic heat calculation method.

4. The method for the analytical forecasting of spacecraft reentry merle of claim 1,

in the second step, the flight characteristic calculation analysis adopts a six-degree-of-freedom or three-degree-of-freedom trajectory calculation method; when the movement details before disassembly need to be mastered, a six-degree-of-freedom ballistic equation can be adopted; the ballistic equation set can be solved step by using a 4-order Runge-Kutta method integral.

5. The method for analyzing and forecasting spacecraft reentry meteorology of claim 1, wherein in step three, for flight environment, aerodynamic force/heat, flight motion and attitude parameters with uncertainty in spacecraft meteorology process simulation modeling analysis, the upper and lower limits, namely conditional boundaries, are determined according to existing basic research result evaluation; constructing a proper parameter distribution mode in the boundary value domains of the conditions during analysis of the meteority process; constructing a disintegration condition criterion of structure disintegration damage; based on a statistical analysis correlation mathematical method, combining flight motion-force-heat-structure destruction disintegration calculation to carry out quantitative analysis on the concerned target parameters; obtaining statistically significant spacecraft flight path, disintegration process, debris viability and ground risk analysis evaluation results in the form of distribution bands with a certain degree of confidence; the parameter distribution refers to each condition parameter with uncertainty, the condition parameters are set to be normal distribution with a reference value as a center, namely Gaussian distribution, the random value range is based on the reference value, the deviation is plus-minus three times of standard deviation, namely the parameter coverage rate is about 99.5% theoretically;

the condition state parameters related to the condition boundary cover performance parameters in the aspects of flight motion, aerodynamic force and aerodynamic heat, and comprise key time points such as reentry time parameters and disintegration time parameters, and physical parameters of a penetration process such as atmospheric density and aerodynamic force coefficients; in the preliminary induction, the statistical analysis parameters related to the spacecraft merle analysis forecasting method comprise: initial position, initial velocity, aerodynamic coefficient, atmospheric density, height of disintegration, property of disintegration fragments, and attitude of disintegration fragments;

the disintegration condition criteria include: height criteria: i.e. a melting criterion is set for a certain height value for some form of disintegration process to occur: namely, the metal material is set to have a certain form of disintegration process at the melting temperature; and (3) pyrolysis criterion: namely, the composite material is set to be subjected to a certain form of disintegration process after being completely pyrolyzed at the pyrolysis temperature; temperature criterion: namely, the object plane temperature is considered to be up to a certain set value and a certain form of disintegration process will occur; heat flow criteria: that is, it is considered that some form of disintegration process will occur when the accumulated heat flow or the instantaneous heat flow reaches a certain set value; dynamic pressure criterion: namely, the dynamic pressure reaches a certain set value and a certain form of disintegration process is generated; and (3) comprehensive criteria: the combination of the two or more criteria can comprise the weight sum, the achievement of any criterion and the achievement of all criteria; the melting criterion and the pyrolysis criterion are a composite form of a temperature criterion and a heat flow criterion; the melting rule is listed to facilitate the application analysis of the metal material with the largest proportion in the current spacecraft structural materials; the pyrolysis criterion is listed to carry out application analysis on the main material composite materials of the container class with special requirements on the current spacecraft.

6. The method for the analytical forecasting of spacecraft reentry merle of claim 1,

in the fifth step, the analysis result of aerodynamic force and aerodynamic heat is used as the boundary condition of the object plane for structural stress analysis; obtaining physical property parameters changing along with temperature and thermal stress caused by temperature difference by thermal analysis in the structure; under the comprehensive action of object surface load transfer stress and thermal stress, analyzing structural damage according to the total stress distribution condition of the material and the Young modulus and Poisson ratio parameters under corresponding conditions;

in the sixth step, the determination of the spacecraft component level disintegration criterion parameter and the conditional boundary parameter is consistent with the determination of the spacecraft system/subsystem level disintegration criterion parameter and the conditional boundary parameter in the third step, and the parameters include: the initial position and the initial speed of the component level and the parameters of the inherited system/subsystem level disintegration time; the rest of the components comprise aerodynamic coefficient, atmospheric density and disintegration height;

in the seventh step, a three-degree-of-freedom ballistic equation and a aerodynamic heat rapid calculation method are adopted for calculating and simulating the component-level flight path; and calculating attenuation change of the mass of the research object according to the fusion ablation rate in the simulation process.

7. The method for analyzing and forecasting reentry of spacecraft into meteor space of claim 1, wherein in the ninth step, geometric features of the spacecraft are highly abstracted in the modeling of the shape at the debris/particle level; the geometric modeling abstraction is based on the principle of a complete induction method, and different geometric characteristics can be covered, so that corresponding pneumatic characteristics and flight motion characteristics are covered and embodied; the specific method is that the fragments/particles are divided into the following parts according to geometrical characteristics: blocks, strips, sheets; this mode of division of the spacecraft debris/particle level is therefore referred to as the patch model.

8. The method for the analytical forecasting of spacecraft reentry merle of claim 7,

in the eleventh step, aerodynamic force and flight path of the strip are calculated and simulated according to the strip model, and the aerodynamic force adopts aerodynamic force coefficient tables of different shapes for the strips to perform interpolation; and calculating and simulating the flight path by adopting a three-degree-of-freedom trajectory equation.

9. The method for the analytical forecasting of spacecraft reentry merle of claim 8,

in the twelfth step, analyzing and evaluating the landing area scattering according to the flight path data of the debris fragments obtained in the eleventh step; the landing zone range depends on the following factors: flight parameters and uncertainty at the time of disassembling the whole device and the parts, and the material, shape, size and distribution rule of debris fragments.

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