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Articles prepublish have been peer-reviewed and accepted, which are not yet assigned to volumes /issues, but are citable by Digital Object Identifier (DOI).
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Multi-satellite precision Orbit Determination and data Analysis Software in solar system (MODAS) is produced by the orbit dynamics group at the Science and Technology on Aerospace Flight Dynamics Laboratory. It provides an efficient and uniformed approach to perform orbit determination for spacecraft in solar system. In addition, it is aimed at solving the joint orbit determination of multiple spacecraft orbiting different celestial bodies, and can be used for both scientific application analysis and engineering practice. This paper describes the design and implementation of MODAS, including data organization and management, fundamental services, measurement models, and parameter estimation. Then, the implementation characteristics of MODAS are introduced in detail. The function is validated through the exploration data processing of the moon, Mars, Jupiter and asteroids. Finally, further application and improvements of MODAS is discussed.
Multi-satellite precision Orbit Determination and data Analysis Software in solar system (MODAS) is produced by the orbit dynamics group at the Science and Technology on Aerospace Flight Dynamics Laboratory. It provides an efficient and uniformed approach to perform orbit determination for spacecraft in solar system. In addition, it is aimed at solving the joint orbit determination of multiple spacecraft orbiting different celestial bodies, and can be used for both scientific application analysis and engineering practice. This paper describes the design and implementation of MODAS, including data organization and management, fundamental services, measurement models, and parameter estimation. Then, the implementation characteristics of MODAS are introduced in detail. The function is validated through the exploration data processing of the moon, Mars, Jupiter and asteroids. Finally, further application and improvements of MODAS is discussed.
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Abstract:
Reusable launch vehicle is one of the important ways to reduce the cost of launch service. This paper focuses on the modeling difficulty problem on the original fatigue load data of reusable launch vehicle engine. In this paper, the root mean square value is selected as the division standard for the original fatigue load data of the reusable launch vehicle. Original data are processed by modified short-time Fourier wave filtering, rain flow cycle counting and Gaussian distribution fitting for the identification and regularization of fatigue load data. Fatigue load data of reusable launch vehicle can be described by Gaussian distribution model. The Gaussian distribution parameter of abnormal fatigue load data is more than 3 times of normal fatigue load data This method can be used to accurately identify the abnormal fatigue load data. Compared with traditional anomaly data identification methods, this method can provide the quantitative index of abnormal data, which provides a new analysis method for fatigue load design and real-time fault analysis and location of reusable launch vehicle.
Reusable launch vehicle is one of the important ways to reduce the cost of launch service. This paper focuses on the modeling difficulty problem on the original fatigue load data of reusable launch vehicle engine. In this paper, the root mean square value is selected as the division standard for the original fatigue load data of the reusable launch vehicle. Original data are processed by modified short-time Fourier wave filtering, rain flow cycle counting and Gaussian distribution fitting for the identification and regularization of fatigue load data. Fatigue load data of reusable launch vehicle can be described by Gaussian distribution model. The Gaussian distribution parameter of abnormal fatigue load data is more than 3 times of normal fatigue load data This method can be used to accurately identify the abnormal fatigue load data. Compared with traditional anomaly data identification methods, this method can provide the quantitative index of abnormal data, which provides a new analysis method for fatigue load design and real-time fault analysis and location of reusable launch vehicle.
, Available online
Abstract:
Aiming at the control difficulties in the recover mission of vertical landing of returning rocket,such as static unstable shape , elastic deformation of large slenderness ratio arrow body, liquid storage phenomenon for zero-pole ,large interference of sudden change of aerodynamic shape during the deployment of landing mechanism,large -scale attitude adjustment,etc.The requirements and constraints of the most critical vertical landing phase on the control system is analyzed comprehensively in this paper. More over, a high-precision attitude control scheme for vertical landing phase is put forward,that include parametric optimization design for large static instability and slosh-polar zero structure, nonlinear smooth gain scheduling ,and high precision attitude control algorithm based on ESO.Simulation results show that the design requirenments can be meet by scheme proposed in this paper .
Aiming at the control difficulties in the recover mission of vertical landing of returning rocket,such as static unstable shape , elastic deformation of large slenderness ratio arrow body, liquid storage phenomenon for zero-pole ,large interference of sudden change of aerodynamic shape during the deployment of landing mechanism,large -scale attitude adjustment,etc.The requirements and constraints of the most critical vertical landing phase on the control system is analyzed comprehensively in this paper. More over, a high-precision attitude control scheme for vertical landing phase is put forward,that include parametric optimization design for large static instability and slosh-polar zero structure, nonlinear smooth gain scheduling ,and high precision attitude control algorithm based on ESO.Simulation results show that the design requirenments can be meet by scheme proposed in this paper .
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220012
Abstract:
The lightweight structure design of the reflector is a difficult problem in the complex optical mechanical system with multi-channel detection, large number of reflectors, strict requirements for volume and quality, cross folding of optical paths in space, short distance of optical components, and easy interference. Either traditional methods have high lightweight rate, but the process is complex, or the lightweight rate is limited. For solving these problems, a new structural design method of special-shaped lightweight aluminum reflectors for space spectrometer is proposed, which includes material selection, configuration design, assembly deformation unloading mode, lightweight design method, surface coating, detection and error compensation, The experimental verification of a space spectrometer shows that the lightweight rate of all aluminum mirrors designed by this method is better than 50%. The method can be applied to a variety of mirror types involving aluminum mirrors, and covers various application scenarios and requirements of optical aperture, optical surface form and spatial configuration.
The lightweight structure design of the reflector is a difficult problem in the complex optical mechanical system with multi-channel detection, large number of reflectors, strict requirements for volume and quality, cross folding of optical paths in space, short distance of optical components, and easy interference. Either traditional methods have high lightweight rate, but the process is complex, or the lightweight rate is limited. For solving these problems, a new structural design method of special-shaped lightweight aluminum reflectors for space spectrometer is proposed, which includes material selection, configuration design, assembly deformation unloading mode, lightweight design method, surface coating, detection and error compensation, The experimental verification of a space spectrometer shows that the lightweight rate of all aluminum mirrors designed by this method is better than 50%. The method can be applied to a variety of mirror types involving aluminum mirrors, and covers various application scenarios and requirements of optical aperture, optical surface form and spatial configuration.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220007
Abstract:
Planning and control for scientific exploration is one of the key technologies for rover teleoperation. Multi-spectral camera that can detect components in coals and rocks on Mars is mounted on the mast of Zhurong Mars Rover. To achieve the goal of accuracy pointing to particular target, an accurate method to calculate angles of mast joints was proposed. The iterative solution method of mast control parameters and the working flow of rover target detection were designed. The error analysis of the method was conducted. Through the practice of China's first Mars exploration mission, the method is satisfied the demands for detection for narrow-field-of-view multi-spectral camera.
Planning and control for scientific exploration is one of the key technologies for rover teleoperation. Multi-spectral camera that can detect components in coals and rocks on Mars is mounted on the mast of Zhurong Mars Rover. To achieve the goal of accuracy pointing to particular target, an accurate method to calculate angles of mast joints was proposed. The iterative solution method of mast control parameters and the working flow of rover target detection were designed. The error analysis of the method was conducted. Through the practice of China's first Mars exploration mission, the method is satisfied the demands for detection for narrow-field-of-view multi-spectral camera.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20210146
Abstract:
This paper proposes a reusable LO/LH propulsion based single-stage manned lunar horizontal attitude vertical landing and ascent spacecraft scheme. The scheme is simple and reliable, capable of matching new-generation human launch vehicles, and enabling reusability via lunar in-situ resource utilization. Meanwhile, a convex optimization based guidance method is proposed for horizontal attitude vertical landing brake phase in view of great fuel consumption and low control accuracy. The method possesses good mission applicability, enabling fuel cost optimization with various constraints including landing attitude and position accuracy. The proposed spacecraft scheme can be applied to lunar and cis-lunar reusable transportation missions, and provides a possible integrated transportation vehicle for both Earth orbit and lunar surface, holding a high cost-effectiveness ratio. The proposed scheme is also provide a good solution for future transportation system enabling large-scale cis-lunar exploration and exploitation.
This paper proposes a reusable LO/LH propulsion based single-stage manned lunar horizontal attitude vertical landing and ascent spacecraft scheme. The scheme is simple and reliable, capable of matching new-generation human launch vehicles, and enabling reusability via lunar in-situ resource utilization. Meanwhile, a convex optimization based guidance method is proposed for horizontal attitude vertical landing brake phase in view of great fuel consumption and low control accuracy. The method possesses good mission applicability, enabling fuel cost optimization with various constraints including landing attitude and position accuracy. The proposed spacecraft scheme can be applied to lunar and cis-lunar reusable transportation missions, and provides a possible integrated transportation vehicle for both Earth orbit and lunar surface, holding a high cost-effectiveness ratio. The proposed scheme is also provide a good solution for future transportation system enabling large-scale cis-lunar exploration and exploitation.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220035
Abstract:
Traditional reliability evaluation of liquid rocket engines requires a large number of verication samples which is too expensive and difficult for engineering application.Combined with the requirements for reusable liquid rocket engines in the context of resuable rocket mission, the reliability evaluation method of reusable liquid rocket engine is studied. By analyzing the characteristics and failure modes of the instantaneous process of ignition for the reusable rocket engines, it is appropriate that regarding the instantaneous process of ignition as two independent processes which contains ignition process and transient development process, and applying the maximum entropy method and normal allowable limit method individually can obtain higher reliability evaluation results with fewer tests. Meanwhile, according to the reliability evaluation method studied, the influence of reliability requirements and life shape parameters to total time of hot-fire tests, as well as the change of reliability with the number of tasks performed have been analyzed, It is shown that the proposed evaluation method can obtain higher reliability evaluation results with less test times.With the higher reliability requirements, the total time of hot-fire tests increases greatly; increasing the life shape parameters can greatly reduce the total time of hot-fire test with the same reliability requirements; and the engine reliability decreases gradually with the increase number of task performed.
Traditional reliability evaluation of liquid rocket engines requires a large number of verication samples which is too expensive and difficult for engineering application.Combined with the requirements for reusable liquid rocket engines in the context of resuable rocket mission, the reliability evaluation method of reusable liquid rocket engine is studied. By analyzing the characteristics and failure modes of the instantaneous process of ignition for the reusable rocket engines, it is appropriate that regarding the instantaneous process of ignition as two independent processes which contains ignition process and transient development process, and applying the maximum entropy method and normal allowable limit method individually can obtain higher reliability evaluation results with fewer tests. Meanwhile, according to the reliability evaluation method studied, the influence of reliability requirements and life shape parameters to total time of hot-fire tests, as well as the change of reliability with the number of tasks performed have been analyzed, It is shown that the proposed evaluation method can obtain higher reliability evaluation results with less test times.With the higher reliability requirements, the total time of hot-fire tests increases greatly; increasing the life shape parameters can greatly reduce the total time of hot-fire test with the same reliability requirements; and the engine reliability decreases gradually with the increase number of task performed.
, Available online
Abstract:
Different from the current reusable launch vehicle landing leg recovery scheme, this paper proposes a rocket vertical landing recovery mechanism scheme based on the ground arresting principle. In the process of rocket vertical landing and descent, the comprehensive action of ground arresting cable and damper is used to realize rocket braking recovery. Through the simulation analysis and the verification of the principle prototype, the ground arresting mechanism can realize the arresting and recovery of the launch vehicle under the conditions of normal landing point, horizontal offset and rotation, the overload and impact borne by the rocket structure are small, and the attitude is stable. The research shows that the arresting recovery mechanism can realize the safe and controllable recovery of the rocket body structure.
Different from the current reusable launch vehicle landing leg recovery scheme, this paper proposes a rocket vertical landing recovery mechanism scheme based on the ground arresting principle. In the process of rocket vertical landing and descent, the comprehensive action of ground arresting cable and damper is used to realize rocket braking recovery. Through the simulation analysis and the verification of the principle prototype, the ground arresting mechanism can realize the arresting and recovery of the launch vehicle under the conditions of normal landing point, horizontal offset and rotation, the overload and impact borne by the rocket structure are small, and the attitude is stable. The research shows that the arresting recovery mechanism can realize the safe and controllable recovery of the rocket body structure.
, Available online
Abstract:
reusable launch vehicle can take off vertically and horizontally, and then land vertically, horizontally, or by parachute, so as to form various combinations of takeoff and landing ways. Aiming at the different recovery modes of reusable launch vehicle, this paper discusses the key technologies of vertical takeoff and vertical landing (VTVL), parachute recovery, and horizontal takeoff and horizontal landing (HTHL), covering the key technologies of other combinations. For VTVL mode, three key technologies, such as engine throttling, multiple start-up, and landing mechanism, are analyzed in detail. For parachute recovery, the technologies relating to the landing area control of rocket jettisons and aerial recovery are introduced. For the HTHL mode based on the rocket propulsion system, the challenges arising from five aspects, including the coupling mechanism under complex aerodynamic thermal environment, thermal protection, landing mechanisms, guidance and control, are discussed. The characteristics of the three recovery modes are briefly summarized and compared with each other in the end.
reusable launch vehicle can take off vertically and horizontally, and then land vertically, horizontally, or by parachute, so as to form various combinations of takeoff and landing ways. Aiming at the different recovery modes of reusable launch vehicle, this paper discusses the key technologies of vertical takeoff and vertical landing (VTVL), parachute recovery, and horizontal takeoff and horizontal landing (HTHL), covering the key technologies of other combinations. For VTVL mode, three key technologies, such as engine throttling, multiple start-up, and landing mechanism, are analyzed in detail. For parachute recovery, the technologies relating to the landing area control of rocket jettisons and aerial recovery are introduced. For the HTHL mode based on the rocket propulsion system, the challenges arising from five aspects, including the coupling mechanism under complex aerodynamic thermal environment, thermal protection, landing mechanisms, guidance and control, are discussed. The characteristics of the three recovery modes are briefly summarized and compared with each other in the end.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20210152
Abstract:
The aerodynamic characteristicsof the rocket first sub-stage with grid rudder and glider was analyzed using 3-D CFD numerical simulation method, taking Falcon-9 and New Glenn rocket as research object. The result show that the axis and normal force of the sub-stage with grid rudder first increased then decreased with the increase of Mach number. For the small angle-of-attack cases, the axis and normal force of the sub-stage with glider decreased with the increase of Mach number when Ma>2. The comparison results show that the static stability of the sub-stage with grid rudder was better than the one with glider, which means that the sub-stage with grid rudder was more suitable for the mission requiring accurate control of the landing point. Meanwhile, the lift drag ratio of the sub-stage with glider is higher than the one with grid rudder, indicating that the sub-stage with glider was more applicable for long distance gliding.
The aerodynamic characteristicsof the rocket first sub-stage with grid rudder and glider was analyzed using 3-D CFD numerical simulation method, taking Falcon-9 and New Glenn rocket as research object. The result show that the axis and normal force of the sub-stage with grid rudder first increased then decreased with the increase of Mach number. For the small angle-of-attack cases, the axis and normal force of the sub-stage with glider decreased with the increase of Mach number when Ma>2. The comparison results show that the static stability of the sub-stage with grid rudder was better than the one with glider, which means that the sub-stage with grid rudder was more suitable for the mission requiring accurate control of the landing point. Meanwhile, the lift drag ratio of the sub-stage with glider is higher than the one with grid rudder, indicating that the sub-stage with glider was more applicable for long distance gliding.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220043
Abstract:
Selecting features for use with SPC navigation requires high-level principal and massive testing to ensure good matching performance. In this work, by analyzing the processes of shape model reconstruction, image rendering, template matching and spacecraft localization, we propose four principles of feature selection applying to natural feature tracking during aircraft descent phase. We validate our principles using real imagery from NASA's OSIRIS-Rex mission to asteroid Bennu, along with the accuracy of local elevations, the correlation performance and the navigation comparison data from NASA NAIF SPICE kernels. The results show that the successful matching rate of rendered features against onboard imagery can reach 96.85% and average correlation score above 0.721 5.
Selecting features for use with SPC navigation requires high-level principal and massive testing to ensure good matching performance. In this work, by analyzing the processes of shape model reconstruction, image rendering, template matching and spacecraft localization, we propose four principles of feature selection applying to natural feature tracking during aircraft descent phase. We validate our principles using real imagery from NASA's OSIRIS-Rex mission to asteroid Bennu, along with the accuracy of local elevations, the correlation performance and the navigation comparison data from NASA NAIF SPICE kernels. The results show that the successful matching rate of rendered features against onboard imagery can reach 96.85% and average correlation score above 0.721 5.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220063
Abstract:
Aiming at the requirements of the landing buffer test system for small space objects in the future, as well as the shortcomings of the existing vertical constant force unloading technology, such as poor stability and low dynamic accuracy. This work proposes a simulation method of heavy load and large travel near zero stiffness support based on the principle of magnetic zero stiffness. The dynamic model and control model of the system are established. And the dynamic characteristics of the system are simulated and analyzed by Simulink. The prototype of the landing buffer system for small celestial bodies is completed, and the landing buffer tests under micro/low gravity conditions are carried out. The results show that the simulation results are in good agreement with the test results. The force resolution of the system is better than 0.1‰ (0.1N), and the constant force error under landing impact condition is less than 0.1%. This system can be applied to the impact test of small celestial landing collision, and greatly improves the test accuracy of small celestial ground simulation test.
Aiming at the requirements of the landing buffer test system for small space objects in the future, as well as the shortcomings of the existing vertical constant force unloading technology, such as poor stability and low dynamic accuracy. This work proposes a simulation method of heavy load and large travel near zero stiffness support based on the principle of magnetic zero stiffness. The dynamic model and control model of the system are established. And the dynamic characteristics of the system are simulated and analyzed by Simulink. The prototype of the landing buffer system for small celestial bodies is completed, and the landing buffer tests under micro/low gravity conditions are carried out. The results show that the simulation results are in good agreement with the test results. The force resolution of the system is better than 0.1‰ (0.1N), and the constant force error under landing impact condition is less than 0.1%. This system can be applied to the impact test of small celestial landing collision, and greatly improves the test accuracy of small celestial ground simulation test.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20210110
Abstract:
Research on the scheme of the vertical landing mechanism of the launch vehicle, with graph theory analysis, topology analysis, the design schemes of different landing mechanisms are analyzed and optimized. For at high reliability, the mechanism scheme of a single closed-loop chain is determined. To the goal of high efficiency, the relevant motion pair selection scheme is determined. Considering the simplification of mechanism energy and other factors, a new design scheme of landing mechanism based on mortise lock is proposed. The research results can provide reference and technical support for the development of the vertical landing mechanism of the reusable launch vehicle.
Research on the scheme of the vertical landing mechanism of the launch vehicle, with graph theory analysis, topology analysis, the design schemes of different landing mechanisms are analyzed and optimized. For at high reliability, the mechanism scheme of a single closed-loop chain is determined. To the goal of high efficiency, the relevant motion pair selection scheme is determined. Considering the simplification of mechanism energy and other factors, a new design scheme of landing mechanism based on mortise lock is proposed. The research results can provide reference and technical support for the development of the vertical landing mechanism of the reusable launch vehicle.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022. 20200074
Abstract:
In this paper, based on the polyhedral model of irregular asteroids, the gravity calculation and position judgment criteria of irregular asteroids are studied and approximated by using Deep neural network（DNN）, and a fast calculation method for the unpowered descent mission of asteroids is proposed. The polyhedron method is used to generate training data, and two DNN models are trained to calculate the gravitational field around the asteroid and judge whether it reaches the asteroid boundary. This method can not only ensure the accuracy of gravity calculation near asteroids, but also save calculation time. Taking Eros 433 asteroid as an example, the trajectory and impact point position of the unpowered descent process are simulated. The simulation results show that the calculation accuracy of the DNN model can meet the mission requirements, the deviation of the impact point position is within a reasonable range, and the calculation efficiency is high, which can be used in large-scale simulation experiments.
In this paper, based on the polyhedral model of irregular asteroids, the gravity calculation and position judgment criteria of irregular asteroids are studied and approximated by using Deep neural network（DNN）, and a fast calculation method for the unpowered descent mission of asteroids is proposed. The polyhedron method is used to generate training data, and two DNN models are trained to calculate the gravitational field around the asteroid and judge whether it reaches the asteroid boundary. This method can not only ensure the accuracy of gravity calculation near asteroids, but also save calculation time. Taking Eros 433 asteroid as an example, the trajectory and impact point position of the unpowered descent process are simulated. The simulation results show that the calculation accuracy of the DNN model can meet the mission requirements, the deviation of the impact point position is within a reasonable range, and the calculation efficiency is high, which can be used in large-scale simulation experiments.
, Available online ,
doi: 10.15982/j.issn.2096-9287.2022.20220045
Abstract:
This paper analyzes and evaluates the navigation and orbit determination performance of the Russian space system "NEBOCVOD", which is designed for monitoring Near Earth Objects (NEOs). And we focus on the optical detectable region and asteroid orbit determination accuracy of the system. Firstly, considering the features of space-based optical measurement system, we propose a method for analyzing the optical detectable region by constructing the contours of apparent magnitude and the apparent motion trajectory of asteroids in the Sun-Earth rotating coordinate system. Then, by modeling the motion and measurement of the system, and considering the influence of uncertain nuisance parameters, we propose a novel approach for orbit determination and accuracy analysis based on improved least square method and covariance analysis. Finally, taking the hazardous NEOs (Apophis, 2008TC3 and Chelyabinsk meteorite) as examples, we simulate the observation and orbit determination process when the "NEBOSVOD" system is assumed to monitor NEOs with different orbital types and sizes, and present the simulation results for multiple measurement intervals. It is shown that the "NEBOCVOD" system can effectively observe medium-sized asteroids like Apophis (D≈320 m) and smaller (D≈4-50 m) NEOs, including those approaching Earth from the Sun, and can achieve high enough navigation and orbit determination accuracy. Our methods and results may provide help for the design and research of the similar space-based monitoring system.
This paper analyzes and evaluates the navigation and orbit determination performance of the Russian space system "NEBOCVOD", which is designed for monitoring Near Earth Objects (NEOs). And we focus on the optical detectable region and asteroid orbit determination accuracy of the system. Firstly, considering the features of space-based optical measurement system, we propose a method for analyzing the optical detectable region by constructing the contours of apparent magnitude and the apparent motion trajectory of asteroids in the Sun-Earth rotating coordinate system. Then, by modeling the motion and measurement of the system, and considering the influence of uncertain nuisance parameters, we propose a novel approach for orbit determination and accuracy analysis based on improved least square method and covariance analysis. Finally, taking the hazardous NEOs (Apophis, 2008TC3 and Chelyabinsk meteorite) as examples, we simulate the observation and orbit determination process when the "NEBOSVOD" system is assumed to monitor NEOs with different orbital types and sizes, and present the simulation results for multiple measurement intervals. It is shown that the "NEBOCVOD" system can effectively observe medium-sized asteroids like Apophis (D≈320 m) and smaller (D≈4-50 m) NEOs, including those approaching Earth from the Sun, and can achieve high enough navigation and orbit determination accuracy. Our methods and results may provide help for the design and research of the similar space-based monitoring system.
Articles just accepted have been peer-reviewed and accepted, which are not yet assigned to volumes /issues, but are citable by Digital Object Identifier (DOI).
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Special Issue: Small Celestial Body Exploration and Defense
(Guest Editor:Professor CUI Pingyuan, Professor, Beijing Institute of Technology, SHANG Haibin, Professor, Beijing Institute of Technology)
2022, 9(4): 359-372.
doi: 10.15982/j.issn.2096-9287.2022.20220074
Abstract:
Small celestial bodies in the Solar system, such as asteroids and comets, have become the key targets in the field of deep space exploration. The exploration of small celestial bodies is of great significance for studying the formation and evolution of the solar system, the origin of life, planetary defense, and the exploitation of space resources. With the continuous development of aerospace technology, the way of small-body exploration has gradually shifted from flyby and orbiting to close-proximity detection methods, such as landing, sampling, and touring. The construction of a high-accuracy gravitational field model in the vicinity of the given small body is crucial to the design and implementation of such close-proximity exploration missions and to the study of dynamics near the small bodies. Thus, this paper first reviews the two-hundred-year history of development of the modeling of gravitational field, and elucidates the basic principles and drawbacks in the different methods. The relationship between the gravitational field near the binary asteroid system and the dynamics of such system is second reviewed. This paper also reviews the techniques for modelling the gravitational interactions in the study of the dynamics of the binary asteroid system. These techniques provide fundamental tools for the binary system mission designs and the study of the formation and evolution of the binaries. Finally, the future research trends are discussed.
Small celestial bodies in the Solar system, such as asteroids and comets, have become the key targets in the field of deep space exploration. The exploration of small celestial bodies is of great significance for studying the formation and evolution of the solar system, the origin of life, planetary defense, and the exploitation of space resources. With the continuous development of aerospace technology, the way of small-body exploration has gradually shifted from flyby and orbiting to close-proximity detection methods, such as landing, sampling, and touring. The construction of a high-accuracy gravitational field model in the vicinity of the given small body is crucial to the design and implementation of such close-proximity exploration missions and to the study of dynamics near the small bodies. Thus, this paper first reviews the two-hundred-year history of development of the modeling of gravitational field, and elucidates the basic principles and drawbacks in the different methods. The relationship between the gravitational field near the binary asteroid system and the dynamics of such system is second reviewed. This paper also reviews the techniques for modelling the gravitational interactions in the study of the dynamics of the binary asteroid system. These techniques provide fundamental tools for the binary system mission designs and the study of the formation and evolution of the binaries. Finally, the future research trends are discussed.
2022, 9(4): 373-381.
doi: 10.15982/j.issn.2096-9287.2022.20210143
Abstract:
Closely flying by asteroids can help to capture asteroid surface images, measure asteroid spectra, and obtain physical and chemical properties of asteroids. In particular, flying by multiple asteroids with potential hazards to the earth in one mission will significantly improve the understanding of the characteristics of potentially hazardous asteroids, and it is also of great significance to asteroid defense missions. In this paper, the trajectory of the multiple asteroid flyby mission of potentially hazardous asteroids was optimized. Firstly, the time and position distribution of asteroids passing through the ecliptic plane were analyzed, and the basic strategy of asteroids flyby in the ecliptic was determined. The time of asteroids crossing the ecliptic was taken as the time of asteroids’ flyby. Secondly, the sequential flyby sequence was optimized via beam selection tree search algorithm, and an optimization model for fast and effective solution of asteroid sequential flyby mission trajectory was established. Simulation results show that missions launched from 2024 to 2028 can fly by at least 18 potentially hazardous asteroids, especially the launch window in September 2027, which can fly by 21 potentially hazardous asteroids within a ten-year mission duration.
Closely flying by asteroids can help to capture asteroid surface images, measure asteroid spectra, and obtain physical and chemical properties of asteroids. In particular, flying by multiple asteroids with potential hazards to the earth in one mission will significantly improve the understanding of the characteristics of potentially hazardous asteroids, and it is also of great significance to asteroid defense missions. In this paper, the trajectory of the multiple asteroid flyby mission of potentially hazardous asteroids was optimized. Firstly, the time and position distribution of asteroids passing through the ecliptic plane were analyzed, and the basic strategy of asteroids flyby in the ecliptic was determined. The time of asteroids crossing the ecliptic was taken as the time of asteroids’ flyby. Secondly, the sequential flyby sequence was optimized via beam selection tree search algorithm, and an optimization model for fast and effective solution of asteroid sequential flyby mission trajectory was established. Simulation results show that missions launched from 2024 to 2028 can fly by at least 18 potentially hazardous asteroids, especially the launch window in September 2027, which can fly by 21 potentially hazardous asteroids within a ten-year mission duration.
2022, 9(4): 382-390.
doi: 10.15982/j.issn.2096-9287.2022.20220024
Abstract:
This paper applied a polyhedron-ellipsoid model to study the dynamics of resonant orbits near the binary system, with the binary system 66391 Moshup as an example. A series of resonant orbital families were calculated the shooting method and continuation method, and the stability and bifurcation of the orbital families were analyzed. Finally, homoclinic connections between resonant orbits were computed by using invariant manifolds. The research shows that no strict planar resonant orbit exists because of the asymmetry of the gravitational field, and the orbital stability and bifurcation have also been affected. In addition, the availability of designing transfer trajectories through resonant orbits was also demonstrated.
This paper applied a polyhedron-ellipsoid model to study the dynamics of resonant orbits near the binary system, with the binary system 66391 Moshup as an example. A series of resonant orbital families were calculated the shooting method and continuation method, and the stability and bifurcation of the orbital families were analyzed. Finally, homoclinic connections between resonant orbits were computed by using invariant manifolds. The research shows that no strict planar resonant orbit exists because of the asymmetry of the gravitational field, and the orbital stability and bifurcation have also been affected. In addition, the availability of designing transfer trajectories through resonant orbits was also demonstrated.
2022, 9(4): 391-399.
doi: 10.15982/j.issn.2096-9287.2022.20220044
Abstract:
During asteroid exploration missions, the spacecraft needs to make an investigation into the characteristics of the asteroid after approaching it. In this paper, the performance of the fly-around mapping orbits and flyby mapping orbits were evaluated. The stable fly-around orbits and hyperbolic flyby orbits at different distances were analyzed using numerical methods. A method to calculate the coverage rate of the orbits was proposed considering the sunlight condition and topographical shading. Then mapping orbits were analyzed in the perspective of coverage rate, mapping time, fuel cost and robustness. Taking nearly spherical asteroid Bennu 101955 and elongated asteroid Eros 433 as examples, the performance of the fly-around orbits and flyby orbits were evaluated, which will provide reference for the mapping phase in future asteroid explorations.
During asteroid exploration missions, the spacecraft needs to make an investigation into the characteristics of the asteroid after approaching it. In this paper, the performance of the fly-around mapping orbits and flyby mapping orbits were evaluated. The stable fly-around orbits and hyperbolic flyby orbits at different distances were analyzed using numerical methods. A method to calculate the coverage rate of the orbits was proposed considering the sunlight condition and topographical shading. Then mapping orbits were analyzed in the perspective of coverage rate, mapping time, fuel cost and robustness. Taking nearly spherical asteroid Bennu 101955 and elongated asteroid Eros 433 as examples, the performance of the fly-around orbits and flyby orbits were evaluated, which will provide reference for the mapping phase in future asteroid explorations.
2022, 9(4): 400-406.
doi: 10.15982/j.issn.2096-9287.2022.20220025
Abstract:
Deep learning algorithm has a higher recognition rate for navigation landmarks such as small meteor craters than traditional algorithms, but it is difficult to achieve matching under various image changes. To solve this problem, a description method of recognition prediction box based on feature descriptor was proposed, and the matching of recognition results was completed. Firstly, the circular support region of the recognition prediction frame was determined and a 10-dimensional feature descriptor with rotation and translation scale and luminance invariance was constructed and the prediction frame was matched by the relative distance between descriptor vectors. The results show that the proposed algorithm is robust to images under different transformations, and the correct matching rate of the prediction frame is over 90%. It may provide the reference for the asteroid exploration navigation system.
Deep learning algorithm has a higher recognition rate for navigation landmarks such as small meteor craters than traditional algorithms, but it is difficult to achieve matching under various image changes. To solve this problem, a description method of recognition prediction box based on feature descriptor was proposed, and the matching of recognition results was completed. Firstly, the circular support region of the recognition prediction frame was determined and a 10-dimensional feature descriptor with rotation and translation scale and luminance invariance was constructed and the prediction frame was matched by the relative distance between descriptor vectors. The results show that the proposed algorithm is robust to images under different transformations, and the correct matching rate of the prediction frame is over 90%. It may provide the reference for the asteroid exploration navigation system.
2022, 9(4): 407-416.
doi: 10.15982/j.issn.2096-9287.2022.20220019
Abstract:
A robust and efficient landmark matching algorithm was proposed in this paper to deal with the extreme environment near the target asteroid. First, the matching error of the landmark generated by the SPC (StereoPhotoClinometry) technology was analyzed, and the influence of landmark position error and cameras’ pose error on the matching results was discussed. Then, based on the error analysis, the optimal landmark points were selected, and a weighted normalized cross-correlation (WNCC) algorithm was proposed to obtain accurate matching results robustly and efficiently. Finally, high-fidelity synthetic image sequences were generated to compare the performance of WNCC and the widely used NCC (Normalized Cross-Correlation) algorithm in the previous asteroid missions under a wide range of image scales, viewing geometries, and lighting conditions. The numerical results demonstrate the advance of the proposed method in terms of efficiency, robustness, and accuracy.
A robust and efficient landmark matching algorithm was proposed in this paper to deal with the extreme environment near the target asteroid. First, the matching error of the landmark generated by the SPC (StereoPhotoClinometry) technology was analyzed, and the influence of landmark position error and cameras’ pose error on the matching results was discussed. Then, based on the error analysis, the optimal landmark points were selected, and a weighted normalized cross-correlation (WNCC) algorithm was proposed to obtain accurate matching results robustly and efficiently. Finally, high-fidelity synthetic image sequences were generated to compare the performance of WNCC and the widely used NCC (Normalized Cross-Correlation) algorithm in the previous asteroid missions under a wide range of image scales, viewing geometries, and lighting conditions. The numerical results demonstrate the advance of the proposed method in terms of efficiency, robustness, and accuracy.
2022, 9(4): 417-426.
doi: 10.15982/j.issn.2096-9287.2022.20220041
Abstract:
Aiming at the requirements of high frame rate, high resolution and high ranging accuracy in small celestial body detection, the characteristics of laser detection technology were deeply analyzed, and a hybrid solid-state laser 3D terrain mapping and navigation integrated design method was proposed. High imaging frame rate was realized by single photon array device and 532nm fiber laser, large field of view and sub-pixel resolution were realized by multi-mode scanning of two-dimensional voice coil motor fast mirror, and high-precision beam expansion and diffraction of laser beam were realized by Damman grating beam splitter. The results show that the laser ranging accuracy is better than 3 cm (3 sigma), the frame rate is 4 Hz, and the imaging resolution is as high as 1 100×1 100. The proposed method can give consideration to both topographic mapping and navigation, realize muti-function, light and miniaturized design, and greatly reduce resource consumption. It has good guiding significance for the implementation of small celestial body exploration missions.
Aiming at the requirements of high frame rate, high resolution and high ranging accuracy in small celestial body detection, the characteristics of laser detection technology were deeply analyzed, and a hybrid solid-state laser 3D terrain mapping and navigation integrated design method was proposed. High imaging frame rate was realized by single photon array device and 532nm fiber laser, large field of view and sub-pixel resolution were realized by multi-mode scanning of two-dimensional voice coil motor fast mirror, and high-precision beam expansion and diffraction of laser beam were realized by Damman grating beam splitter. The results show that the laser ranging accuracy is better than 3 cm (3 sigma), the frame rate is 4 Hz, and the imaging resolution is as high as 1 100×1 100. The proposed method can give consideration to both topographic mapping and navigation, realize muti-function, light and miniaturized design, and greatly reduce resource consumption. It has good guiding significance for the implementation of small celestial body exploration missions.
2022, 9(4): 427-437.
doi: 10.15982/j.issn.2096-9287.2022.20220010
Abstract:
To solve the problem that it is difficult to carry out experimental verification for optical image navigation of small celestial body detector landing section, a small celestial body terrain dynamic simulation and landing scene simulation system was designed based on virtual reality, and the detector attitude design and landing sequence image acquisition were realized. The small celestial body model was established by using three-dimensional modeling technology, and the weighted least square method was used to realize the smooth grid of the model; linear interpolation was used to process the map, and texture mapping was realized by combining spherical mapping with cube mapping; the virtual scene simulation was developed, and the detector attitude design was realized according to the rotation matrix method. Experiments show that the simulation system can meet the demand for image navigation verification of the landing section of the probe, and can realize dynamic observation of small object topography and landing image acquisition with high graphic quality and good real-time performance, and the effectiveness of the system is verified by specific example simulation.
To solve the problem that it is difficult to carry out experimental verification for optical image navigation of small celestial body detector landing section, a small celestial body terrain dynamic simulation and landing scene simulation system was designed based on virtual reality, and the detector attitude design and landing sequence image acquisition were realized. The small celestial body model was established by using three-dimensional modeling technology, and the weighted least square method was used to realize the smooth grid of the model; linear interpolation was used to process the map, and texture mapping was realized by combining spherical mapping with cube mapping; the virtual scene simulation was developed, and the detector attitude design was realized according to the rotation matrix method. Experiments show that the simulation system can meet the demand for image navigation verification of the landing section of the probe, and can realize dynamic observation of small object topography and landing image acquisition with high graphic quality and good real-time performance, and the effectiveness of the system is verified by specific example simulation.
2022, 9(4): 438-446.
doi: 10.15982/j.issn.2096-9287.2022.20220054
Abstract:
According to the requirements of near-Earth Asteroid kinetic impact mission, an autonomous GNC scheme is proposed, which contains the high-precision extraction method for the line of sight （LOS） of the asteroid center, high-precision autonomous relative navigation based on LOS measurement of the asteroid center, and iterative prediction guidance method. In this scheme, the LOS measurement of the asteroid center is used to estimate the relative position and velocity between the impactor and the asteroid in the direction perpendicular to the LOS. Although the elative position and velocity errors along the LOS are not estimated, the high-precision impact can be realized by the iterative prediction guidance method. The mathematical simulation shows that the proposed scheme can ensure the impactor to hit a near-Earth asteroid with the diameter 50 meter, the impact accuracy is better than 4 meter, which meets the mission requirements.
According to the requirements of near-Earth Asteroid kinetic impact mission, an autonomous GNC scheme is proposed, which contains the high-precision extraction method for the line of sight （LOS） of the asteroid center, high-precision autonomous relative navigation based on LOS measurement of the asteroid center, and iterative prediction guidance method. In this scheme, the LOS measurement of the asteroid center is used to estimate the relative position and velocity between the impactor and the asteroid in the direction perpendicular to the LOS. Although the elative position and velocity errors along the LOS are not estimated, the high-precision impact can be realized by the iterative prediction guidance method. The mathematical simulation shows that the proposed scheme can ensure the impactor to hit a near-Earth asteroid with the diameter 50 meter, the impact accuracy is better than 4 meter, which meets the mission requirements.
2022, 9(4): 447-454.
doi: 10.15982/j.issn.2096-9287.2022.20220042
Abstract:
To address the problem that wheeled mobile rovers are difficult to adapt to the weak gravitational environment of small bodies. The jumping mode of the rover with cubic configuration was analyzed. A single-step bouncing strategy was proposed to exploit its barrier-crossing capability. At the same time, a variable-step A* algorithm was proposed to plan the surface movement path of the rover. The simulation results show that the rover can effectively cross the obstacles. The new algorithm reduces the path nodes and is more efficient for high-density obstacle terrain.
To address the problem that wheeled mobile rovers are difficult to adapt to the weak gravitational environment of small bodies. The jumping mode of the rover with cubic configuration was analyzed. A single-step bouncing strategy was proposed to exploit its barrier-crossing capability. At the same time, a variable-step A* algorithm was proposed to plan the surface movement path of the rover. The simulation results show that the rover can effectively cross the obstacles. The new algorithm reduces the path nodes and is more efficient for high-density obstacle terrain.
Founded in 2014, Bimonthly
Supervisor: Ministry of Industry and Information Technology
Sponsor: Beijing Institute of Technology, China Aerospace Society Committee for Deep Space Detection Technology
Editor-in-chief: Wu WeiRen
ISSN 2096-9287CN 10-1707/V
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Telphone: （010）68915831 / 68915834
E-mail: jdse@bit.edu.cn
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