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25 August 2024, Volume 44 Issue 4 Previous Issue    Next Issue

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Multi-impulse pursuit-evasion game in GEO based on improved dung beetle optimization GUO Yanning, LI Gaojian, YU Yongbin 2024, 44 (4):  1-10.  doi: 10.16708/j.cnki.1000-758X.2024.0052 Abstract ( 583 )   PDF (4364KB) ( 481 )   Save This paper investigates the pursuit-evasion game of GEO with J2 perturbation and impulsive thrust considering perception delay.An optimization model of orbital pursuit strategy is established，considering fuel consumption，single impulse velocity increment，impulse interval time，mission duration，impulse quantity，and terminal distance.The design variables include the number of impulses，the sequence of maneuver moments，and the sequence of impulse increments.The pursuing spacecraft pursues the target spacecraft through multiple impulses.To enhance problem-solving efficiency，an improved Bernoulli dung beetle optimization algorithm（IBDBO）utilizing Bernoulli chaotic mapping and optimal value guidance is proposed.Additionally，Lambert maneuver correction is introduced to address terminal constraint satisfaction issues.The comparison experiments with other intelligent algorithms verify the superiority of this algorithm in terms of convergence speed，convergence stability and optimization efficiency.Furthermore，simulations in real scenarios with perceptual delay demonstrate the effectiveness of this algorithm for planning pursuit strategies.Finally，the causal relation between terminal distance of both sides in the game and the target spacecraft′s maneuvering capabilities，perception delay time is explored. Related Articles | Metrics

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Fuzzy gain-adapting coupling attitude control for under-actuated spacecraft MENG Zhongjie, LU Junjie 2024, 44 (4):  11-19.  doi: 10.16708/j.cnki.1000-758X.2024.0053 Abstract ( 288 )   PDF (3949KB) ( 273 )   Save During rapid orbital maneuvers，aiming at the strong attitude disturbance problem caused by thrust misalignment and installation errors under solid propulsion，an underactuated spacecraft intelligent attitude control method based on thrust vector control technology is proposed.Firstly，the dynamic model of spacecraft attitude error is established，and the underactuated characteristic of thrust vector control inputs are analyzed.Then，considering the issues of strong disturbance uncertainty and weak coupling in the roll channel，an underactuated intelligent control law based on enhanced coupling strategy and adaptive fuzzy observer is designed.The fuzzy logic function is used to approximate the strong disturbance uncertainty term and introduced into the control law to achieve underactuated intelligent attitude control of spacecraft.The stability of the system has been proven through Lyapunov theory.Finally，through numerical simulation and comparison with the hierarchical sliding mode control method，the simulation results show that the designed method can shorten the three-axis attitude stability time by 14%，and can effectively eliminate the static error caused by weak roll channel，which provides a foundation for strong disturbance suppression technology during rapid orbital maneuvers. Related Articles | Metrics

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Precise starlight refraction navigation observation model based on LS-SVM YAN Xu, WANG Dingjie, ZHANG Hongbo, YANG Hang, BAO Weimin 2024, 44 (4):  20-28.  doi: 10.16708/j.cnki.1000-758X.2024.0054 Abstract ( 230 )   PDF (4119KB) ( 280 )   Save The precise observation model for stellar refraction navigation using the least squares support vector machine（LS-SVM）is introduced.The process involves high fidelity simulations of stellar refraction via ray tracing，considering changes of atmospheric parameters with altitude，to determine the nonlinear mapping between true altitude and refraction angle.The LS-SVM algorithm is applied to fit this relationship accurately，creating a precise model for refracted apparent height.Simulations show the model achieves high precision，with an average absolute error of 0.986m.When integrated into a stellar navigation system，it significantly enhances navigation accuracy，reducing the average positioning error to 130.7m and velocity error to 0.1479m/s，thereby validating the effectiveness of the modeling approach.This approach is significant for the high-precision study of autonomous navigation using stellar refraction. Related Articles | Metrics

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A guidance strategy for rendezvous and docking to the space station in the Earth-Moon NRHO orbit XIE Yongchun, CHEN Changqing, LI Xiangyu, LI Zhenyu 2024, 44 (4):  29-39.  doi: 10.16708/j.cnki.1000-758X.2024.0055 Abstract ( 144 )   PDF (2596KB) ( 117 )   Save With the development of space technology，it is possible to build a space station in Earth-Moon space as a transit for Earth-Moon round-trip and entering in the deep space.Rendezvous and docking is one of the key technologies for building an Earth-Moon space station.A guidance strategy for rendezvous and docking from the Earth orbit to the space station in the Earth-Moon NRHO orbit is proposed in this paper，which is suitable for engineering applications.Firstly，the rendezvous and docking process is divided into three sections，i.e.，the large-range orbit transfer section，far-range guidance section，and close-range approaching section.The suitable terminal of large-range orbit transfer is selected according to the eigenvalue of NRHO orbit state transition matrix.The two-impulse guidance method based on the relative motion equation in the three-body problem is adopted for the far-range guidance section.The impulse time and amplitude are solved with the optimization algorithm.The linear constant three-body relative motion equation is proposed for the close-range approaching section，and the rendezvous and docking is completed by a two-stage linear approximation.Finally，a simulation analysis is carried out，and the simulation results show that the adopted dynamics equations and the designed guidance law are effective，and the three flight phases are naturally connected to accomplish the rendezvous and docking mission from the Earth orbit to the space station on the Earth-Moon NRHO. Related Articles | Metrics

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Unified numerical predictor-corrector guidance based on characteristic model MENG Bin, ZHANG Hangning, ZHAO Yunbo 2024, 44 (4):  40-49.  doi: 10.16708/j.cnki.1000-758X.2024.0056 Abstract ( 112 )   PDF (455KB) ( 93 )   Save Aerocapture is one of the key technologies for low-cost transportation，with high demands of autonomy，accuracy，and robustness of guidance and control，due to its high reliability requirements for only one chance of trying.A unified numerical predictor-corrector guidance method based on characteristic models for aerocapture is proposed.The numerical predictor-corrector guidance method is used to achieve autonomy and high accuracy，and the characteristic model control method is introduced to achieve robustness.At the same time，by transforming path constraints，characteristic model equations including apogee deviation and altitude differentiation are established.Based on the characteristic model equations，a unified guidance law which can satisfy path constraints and guidance objectives simultaneously is designed.In guidance problems，guidance deviation is not directly obtained from the output of the dynamics at present，but is calculated through integral and algebraic equations.Therefore，the method of directly discretizing differential equations cannot be used to establish characteristic models，which brings great difficulty to characteristic modeling.A method for characteristic modeling of guidance problems is proposed，and convergence analysis of the proposed guidance law is also provided.Finally，a joint numerical simulation of guidance and control considering navigation deviation and various uncertainties is conducted to verify the effectiveness of the proposed method.The proposed unified method can be extended to general aerodynamic entry guidance designs，providing theoretical and methodological support for them. Related Articles | Metrics

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Autonomous capturing and relative navigation methods for small celestial bodies LIU Yiwu, HU Jinchang, LIANG Xiao, TIAN Qihang, ZHANG Hui, YIN Fang 2024, 44 (4):  50-58.  doi: 10.16708/j.cnki.1000-758X.2024.0057 Abstract ( 238 )   PDF (5363KB) ( 268 )   Save Long-range autonomous capture and close-range high-precision autonomous relative navigation are key issues in small celestial body exploration tasks，and the smaller the target size，the more prominent the problem is.Firstly，to deal with the difficulty in identifying dim targets in the starry sky background within a long distance range from tens of thousands to thousands of kilometers，a fully autonomous capture and recognition method that comprehensively utilizes kinematics and brightness is proposed，which can achieve fast and accurate capture of large range changes，and has the ability to capture Mv10 dim targets in the starry sky at a distance of 30000 kilometers.Secondly，to cope with the problem of insufficient observability in line-of-sight measurement in the range of thousands to tens of kilometers during rendezvous，a relative navigation method based on integrated design of line-of-sight and trajectory maneuvering is proposed，which effectively improves trajectory observability while considering fuel consumption，and achieves relative position navigation with an accuracy superior to 5% in the rendezvous segment.Finally，to address the relative navigation problem near irregular small celestial bodies in close range，an optical navigation method based on the combination of image landmarks and point cloud features is proposed，which can simultaneously adapt to the detection needs of sunny and shaded areas.Combined with the weak gravitational force field measuring method based on multi-source data fusion，the accuracy of close-range relative navigation is further improved.The relative position accuracy at close-range is better than 1m，while the relative velocity is better than 1cm/s.The proposed method was validated through mathematical and physical simulations.The method proposed in this article effectively solves the high-precision navigation problem in small celestial body exploration，and can be applied to small celestial body landing detection tasks with diameters of tens of meters and rapid alternation of light and shadow areas. Related Articles | Metrics

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Key technologies and approaches of intelligent control system for flexible wing structure aircraft JIA He, LIU Jinglei, MA Keyao, YAN Yunlong, PEI Xiaoyan, LIU Haiye, WANG Yongbin 2024, 44 (4):  59-70.  doi: 10.16708/j.cnki.1000-758X.2024.0058 Abstract ( 164 )   PDF (2366KB) ( 201 )   Save With the advantages of lightweight，space integration，re-usability，large drag area and lift coefficient，the flexible wing structure aircraft can achieve functions such as cruise flight，low-speed hovering，re-entry and return，aerodynamic deceleration，fixed point homing and landing buffering，which is currently a research hot spot.Intelligent control system is one of the core technologies of the flight and recovery system of flexible wing structure aircraft.Combined with the application research and engineering practice of intelligent technology in the control system，the intelligent control system and its technical characteristics of the flexible wing structure aircraft are analyzed.The key technologies such as integrated control and simulation of rigid-flexible combination，environmental perception and online health status assessment，trajectory planning and tracking control，cluster flight control，intelligent control of landing and buffering，fault tolerance and reconstruction of intelligent hardware are introduced.The future development of the intelligent control system of flexible wing structure aircraft is considered.Some development suggestions are put forward，such as intelligent and flexible perception of flight environment，online identification of aerodynamic parameters，autonomous execution of multi-task mode and evolutionary learning of the control system.Through continuous research and practice of intelligent control technology，strong support is provided for the development of flexible wing structure aircraft system. Related Articles | Metrics

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Review on spacecraft autonomous decision-making and planning for orbital threat avoidance GAO Wanying, WU Jianfa, WEI Chunling 2024, 44 (4):  71-89.  doi: 10.16708/j.cnki.1000-758X.2024.0059 Abstract ( 331 )   PDF (2018KB) ( 661 )   Save The accumulation of space debris，the deployment of large-scale satellite constellations，and intensified spatial competition have led to a rapid increase in the number of orbital threats，seriously threatening the safety and stability of spacecraft.Research into spacecraft autonomous decision-making and planning for orbital threats is crucial to securing China's space assets.Confronted with complex scenarios characterized by high dynamics，time-varying constraints，incomplete or imperfect information，and multiple concurrent threats，this research faces several practical challenges.This review examines the research status of spacecraft autonomous decision-making and planning，discusses key technologies including problem modeling，decision-making，maneuver planning，intelligent decision-making and planning，and concludes with suggestions for future research. Related Articles | Metrics

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Equilibrium of orbital pursuit-evasion-defense three-sided game LI Zhenyu, LIN Kunpeng, HOU Yuzhuo, LUO Yazhong 2024, 44 (4):  90-101.  doi: 10.16708/j.cnki.1000-758X.2024.0060 Abstract ( 295 )   PDF (7093KB) ( 439 )   Save To improve the defense ability of the spacecraft in orbit，an orbital pursuit-evasion-defense（PED）linear-quadratic game is investigated.Three players are called the pursuer，evader，and defender，respectively.The pursuer aims to intercept the evader，while the evader tries to escape from the pursuer，accompanied by a defender who attempts to protect the evader by intercepting the pursuer actively.Due to the existence of the defender，the pursuer has to evade the defender when chasing the evader.Meanwhile，cooperation between the evader and the defender may decrease the difficulty of escape.For such a three-sided game，a linear-quadratic differential game model is established with a performance index combining three players′ energy consumption and the distance.Then the necessary conditions for the Nash equilibrium of the three players are derived and the optimal pursuit guidance law and evasion-defense guidance law are obtained.Furthermore，the equilibrium solution is extended to a more general PED scenario with multiple defenders.Simulation results show that a defender can improve the survivability of the evader.Even with inferiority in maneuverability，they can win the pursuer cooperatively.Besides，an initial position close to the pursuer or evader is not the best choice for the defender who flies around the evader.The defender has favorable positions. Related Articles | Metrics

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Orbit optimization for an asteroid-defense spacecraft departing from large elliptical parking orbit WANG Jie, ZHONG Zikai, YUAN Hao, SONG Haibo 2024, 44 (4):  102-110.  doi: 10.16708/j.cnki.1000-758X.2024.0061 Abstract ( 151 )   PDF (4665KB) ( 211 )   Save Kinetic impact is a means to deflect near-Earth asteroids（NEAs）and to reduce the risk of Earth being impacted by asteroids.An orbit optimization method for spacecraft departing from a large elliptical Earth parking orbit is proposed.Based on genetic algorithm，an optimization model is constructed，which takes the maximum deflection distance of the asteroid as the objective and considers nonlinear constraints on transfer time and fuel of spacecraft.Taking the asteroid 2019 PDC as the object，the genetic algorithm is used to optimize the global optimal solution of the departure date and transfer time of the spacecraft，and to calculate parameters of the large elliptical Earth parking orbit，Earth escape orbit and interplanetary transfer orbit.Meanwhile，Pork-chop plot of the asteroid deflection distance to departure date and transfer time is obtained using the traversing method，which verifies the effectiveness of the optimization algorithm.Results show that the proposed kinetic impact scenario and orbit optimization method can effectively deal with asteroids with diameters of 100 meters and a warning time of several years.The orbit design method and defense scenario proposed can provide a reference for NEAs defense，and lay a foundation for further precise design of spacecraft orbit parameters and implementation of engineering in-orbit verification. Related Articles | Metrics

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Dynamics of asteroid probe orbit calculation WANG Hong, YAN Jianguo, FAN Min, LI Haitao 2024, 44 (4):  111-118.  doi: 10.16708/j.cnki.1000-758X.2024.0062 Abstract ( 223 )   PDF (1959KB) ( 322 )   Save The problem of orbit calculation dynamics in the upcoming asteroid exploration missions of China is studied.The demand of asteroid probe orbit calculation is analyzed，and the space-time reference system involved in asteroid orbit calculation is described，then the dynamical models to be considered in orbit calculation are discussed.The calculation formulas of various types of perturbation acceleration are given.The largest force model is the gravitational acceleration of the Sun，with magnitude about 10-6km/s2，and the smallest is the asteroid body perturbation acceleration which is about 10-12-10-13km/s2.The difference between ephemeris of Rosetta provided by ESA and that integrated from the dynamical model in this paper shows that the four-day arc segment position error is about 15m，velocity error is about 0.44mm/s，so the accuracy of the dynamical model is reliable.The above study can provide a reference for the orbit calculation of China′s asteroid exploration missions. Related Articles | Metrics

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Design of flexible docking mechanism for micro-nano satellites TAO Xinyong, KANG Guohua, WU Junfeng, HU Miaomiao, YANG Zhenghao, ZHOU Shaohui 2024, 44 (4):  119-129.  doi: 10.16708/j.cnki.1000-758X.2024.0063 Abstract ( 192 )   PDF (8096KB) ( 355 )   Save In order to combine micro-nano satellites into ultra-large variable structure spacecraft，the design of relative attitude repeatedly adjustable docking mechanism with high precision is the key problem.Based on Stewart mechanism，a kind of high-precision relative attitude repeatedly adjustable docking mechanism was designed.Aiming at the problem of relative attitude high-precision adjustment after docking locking，the transfer mechanism of attitude adjustment error of docking mechanism was analyzed，and the transfer model of attitude adjustment error was established.The theoretical attitude adjustment accuracy was obtained through simulation calculation.By forming a closed-loop control system with the laser sensor and Stewart mechanism，the error compensation algorithm of the pose control system was designed，the high-precision adjustment of the relative pose was realized.The dynamic model was established in the scenario of two-module relative attitude adjustment，and the driving force variation law of each telescopic rod motor was analyzed according to the motion trajectory of attitude adjustment.Simulation analysis and real object attitude adjustment test show that the weight of the docking mechanism is only 2.43kg，the radius of the envelope is only 10cm，the accuracy of the three-axis attitude adjustment can reach ±0.02°，and the maximum attitude adjustment load of the docking mechanism is 65N，which can be applied to the precise allostery of micro-nano assembly spacecraft. Related Articles | Metrics

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A deep neural network assisted trajectory optimization algorithm for vertical landing vehicles WANG Yazhou, DIAN Songyi, XIANG Guofei 2024, 44 (4):  130-141.  doi: 10.16708/j.cnki.1000-758X.2024.0064 Abstract ( 137 )   PDF (6081KB) ( 203 )   Save It is challenging to solve the powered descent guidance problem online for its computational cost and uncertain initial conditions.An Hp-pseudospectral convex optimization algorithm assisted by deep neural network is presented.For the highly nonlinear dynamics in atmosphere，it is proved for the first time that the thrust magnitude profile has the Bang-Bang feature based on variational method and Pontryagin′s maximum principle.In the wide range of initial states，the deep neural network is applied to learn the segment feature of optimal thrust offline.Then the trained neural net is embedded in the online successive convex optimization algorithm，which combines the Hp-pseudospectral discretization with Bang-Bang feature.This learning assisted strategy leads to more accurate results with the same number of discretized nodes.Numerical simulations show that the proposed algorithm shows better computational efficiency and adaptability to initial conditions. Related Articles | Metrics

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Thermal test of Martian environment based on thermal model correlation verification ZHANG Bingqiang, XUE Shuyan, ZHOU Xiaozhou 2024, 44 (4):  142-152.  doi: 10.16708/j.cnki.1000-758X.2024.0065 Abstract ( 130 )   PDF (8261KB) ( 211 )   Save Considering the difficulty in comprehensively simulating the complex and changeable thermal environment of the Martian surface with a thermal balance test，a thermal balance test method based on the correlation verification of the thermal model is established.In addition，vacuum，low-pressure variable-temperature natural convection and low-pressure constant-temperature forced convection are combined to verify the convection，heat conduction，radiation network coefficient and thermal capacity of the analysis model under high- and low-temperature conditions，and the correctness of the design is verified using the thermal analysis model after correlation verification.The results show that the natural convection heat transfer coefficient under the condition that the plate temperature is 11.86℃ higher than the air temperature（-91℃） was 0.19-0.57W/（m2·K），and the forced convection heat transfer coefficient under the wind speed of 4.15m/s is 1.24-2.4W/（m2·K）.The deviation between the modified thermal analysis model and the test results is within ±3.5℃.This method realizes the thermal design verification of spacecraft under incompletely simulating the thermal environment，provides an effective thermal analysis model，and has been successfully applied to the thermal balance test of the Zhu Rong Mars rover. Related Articles | Metrics

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Research on ultra-wideband high power amplifier technology LI Donglin, YANG Fei 2024, 44 (4):  153-160.  doi: 10.16708/j.cnki.1000-758X.2024.0066 Abstract ( 206 )   PDF (8420KB) ( 338 )   Save Ultra-wideband power amplifier is the core，versatile and high value-added components in wideband communications，electronic countermeasures and radar systems.As the demand for communication systems continues to increase，the need for power amplifiers with wider bandwidth，higher efficiency，and more power is imminent.An ultra-wideband power amplifier is designed based on GaN（gallium nitride）high electron mobility transistor（HEMT）with operating frequency from 0.46GHz to 3.9GHz.Firstly，the input matching is optimized to achieve absolute stability and high gain in the full frequency band.Secondly，the high power and high efficiency output matching impedance space is extracted by load-pull design at different frequency points within the target frequency band，and then a four-stage Chebyshev impedance converter is applied as the output matching topology to achieve high efficiency and high output power in the wide frequency band.The test results show that within the operating band from 0.46GHz to 3.9GHz，the overall output power（Pout）is greater than 13.8W with the maximum output power of 24.1W；the full-band drain efficiency is greater than 50.1% with the maximum value of 67.2% at 1.5GHz.The measured results are in good agreement with the simulation results，and the design idea is intuitive and clear，which provides a direction for designing the currently required ultra-wideband power amplifier.  Related Articles | Metrics

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Research on added mass of disk-gap-band parachute JIANG Tian, GE Sicheng, WANG Yihang 2024, 44 (4):  161-172.  doi: 10.16708/j.cnki.1000-758X.2024.0067 Abstract ( 82 )   PDF (9480KB) ( 103 )   Save The conventional approach to quantifying the added mass of a parachute relies on the premise of an optimal canopy configuration，necessitating the utilization of empirical formulas and coefficients for analysis.Nevertheless，the procedure to identify the added mass is not addressed.A joint simulation scheme is established to facilitate fluid-solid interaction simulation and the calculation of added mass for the disk-gap-band parachute.Correspondingly，a canopy reconstruction strategy is provided for determining the shape of the disk-gap-band parachute and a novel methodology is proposed for the numerical computation of the added mass associated with intricate geometries. The results indicate that through the range of dimensionless inflation time from 0 to 0.6，the increase in added mass of a single disk is equivalent to the aggregate of the parachute.This observation aligns with the mathematical model obtained from established empirical formulas. Following this，a model of a cup with gap is provided that accurately represents the added mass of parachutes at specific time intervals.Correlations are established between the gap structure and other factors such as the height-to-radius ratio，with the aim of deriving a modified empirical formula for predicting the added mass in disk-gap-band parachutes under diverse operational circumstances.The results show that the modified formula is helpful to improve the calculation accuracy of the dynamics simulation in supersonic conditions during the disk-gap-band parachute inflation stage.The research can provide support and reference for the high-precision modeling and analysis for the inflation process of the disk-gap-band parachute. Related Articles | Metrics

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Baseline deviation search algorithm for inter-satellite measurement DAI Yangfeng, LI Mingcheng, WANG Chunhui 2024, 44 (4):  173-183.  doi: 10.16708/j.cnki.1000-758X.2024.0068 Abstract ( 127 )   PDF (8079KB) ( 153 )   Save The baseline deviation introduces significant errors in the process of converting the relative angle measurement results to the main star coordinates.To solve this problem，the conversion deviation introduced by the baseline deviation in the range of azimuth and elevation angles（-60°，+60°）was studied，and a search algorithm for on-orbit calibration of the measurement baseline was proposed.This method used the close-range high-precision perception system carried by the satellite，and at the same time performed perception measurement of the slave star’s position，took the close-range perception system as the calibration data source，then searched for the baseline deviation and calibrated through the step acceleration algorithm，so that the perception results of the system were synchronized with the calibration data source，so as to achieve the purpose of measuring baseline calibration.After the search algorithm，the conversion deviation of azimuth angle and elevation angle is better than 001°，the maximum is 0.012°，which is lower than the maximum measurement error caused by thermal noise. Related Articles | Metrics