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    25 December 2021, Volume 41 Issue 6 Previous Issue   
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    Review and prospect of intelligent perception for non-cooperative targets
    MU Jinzhen, HAO Xiaolong, ZHU Wenshan, LI Shuang
    2021, 41 (6):  1-16.  doi: 10.16708/j.cnki.1000-758X.2021.0076
    Abstract ( 600 )   PDF (1102KB) ( 1381 )   Save
    Intelligent perception is the key technology to realizing the fine control process of spacecraft on-orbit, which is one of the major development directions of on-orbit intelligence service. The key technologies of space target intelligent perception include pose measurement, three-dimensional reconstruction and component recognition, which involve the issues such as few-shot, multi-modality, model adaptation and high-dimensional data. From the perspective of engineering application, the latest research progress in the non-cooperative intelligent perception technique was systematically summarized. Firstly, the research status of the representative non-cooperative on-orbit perception systems and optical sensors were reviewed. Secondly, the key technologies involved in the intelligence perception of non-cooperative targets were analyzed. Finally, according to the analysis of the research status and critical techniques, the main issues of non-cooperative target intelligent perception were discussed, and the recommendations for the further development were presented.
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    Analysis of the law of LEO satellite spin motion under the influence of gravity gradient torque
    CAI Lifeng, ZHANG Guoyun, HONG Tao, LI Weiping, LIN Haichen, SUN Zhenjiang
    2021, 41 (6):  17-24.  doi: 10.16708/j.cnki.1000-758X.2021.0077
    Abstract ( 162 )   PDF (2210KB) ( 185 )   Save
    In order to study the law of LEO satellite spin motion, for the LEO satellite whose ascending node right ascension precesses slowly, its attitude motion model under the influence of gravity gradient torque was established. Besides, the analytical solutions of the precession angle, nutation angle and spin angle were derived when the spin angular rate met certain conditions. Also, the correctness of the analytical solutions was verified by simulation results. In the circumstance that the orbital plane precesses slowly, when the satellite spins around the maximum principal inertia axis, the value range of the spin angular rate was given, and within this range, the spin motion of the satellites can precess with the orbital plane simultaneously, the spin axis can precess at a constant average angular rate, and the nutation angle fluctuates on small scales. When the attitudes of the nearEarth satellites are out of control, the established model on the attitude motion of spinning as well as the analysis conclusions can be used to estimate and predict the attitude. They can also be used for the design of ontrack attitudes as well as ontrack backups.
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    Analysis of influence of solar radiation on communication links during Mars orbiter operation and link parameter design
    SUN Wen, YAN Yi, FAN Yanan, YAO Xiujuan, GAO Xiang, YAN Wenkang
    2021, 41 (6):  25-33.  doi: 10.16708/j.cnki.1000-758X.2021.0078
    Abstract ( 141 )   PDF (4258KB) ( 160 )   Save
    In the super solar junction (SSC) stage of deep space communications, strong solar radiation is one of the key factors that affect the calculation and analysis of the received noise temperature of the ground station. Aiming at the problem of high transmission error rate and communication interruption caused by solar radiation, taking Mars exploration as an example, a research method for calculating Sun-Earth-Probe angle ∠SEP during the operation of the orbiter was proposed. Combined with ∠SEP, the cause for the influence of the Sun on the link in the SSC stage was analyzed. The focus was on the relationship between the solar noise temperature received by the ground station and ∠SEP, air interface parameters, and antenna beam characteristics. Research and simulation experiments show that when the antenna diameter is constant, the peak value of the noise temperature received by the ground station and the influence time of the Sun on the ground station are inversely proportional to the communication frequency. And when the communication frequency is constant, the peak value of the noise temperature received by the ground station is proportional to the antenna diameter. But the influence time of the Sun on the ground station is inversely proportional to the antenna aperture. Combined with the analysis of the influence of solar radiation on the communication link, the link parameter design suggestions under different ∠SEP were given, which provided a reference for the parameter setting in the Mars exploration project.
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    Accuracy analysis of GNSS real-time kinematic timing
    DONG Xiaosong, SUN Baoqi, YANG Haiyan, HAN Baomin, WU Meifang, MENG Lingda, YANG Xuhai
    2021, 41 (6):  34-41.  doi: 10.16708/j.cnki.1000-758X.2021.0079
    Abstract ( 225 )   PDF (3464KB) ( 227 )   Save
    Compared with PPP timing, real-time kinematic timing based on GNSS carrier phase observations can effectively avoid dependence on real-time precise orbit and clock products, and is of great significance to short-distance kinematic and static high-precision time users. In order to verify the performance of GNSS realtime kinematic timing, a timing experiment was carried out with the GPS system as an example based on the time and frequency resources as well as the observation data of three GNSS tracking stations for up to 2 months of the National Time Service Center of the Chinese Academy of Sciences. Compared with the PPP time transfer, the standard deviation of differences from realtime kinematic timing result is better than 0.15ns; compared with the result of optical fiber two-way time transfer, the standard deviation of differences from real-time kinematic timing result is better than 0.5ns. The experiments show that GNSS realtime kinematic timing accuracy can reach the sub-nanosecond level, which can provide an important reference for the future application of this timing technique.
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    Research on tethered space-tug system for non-cooperative spacecraft capture
    SUN Ruiqi, QI Rui
    2021, 41 (6):  42-53.  doi: 10.16708/j.cnki.1000-758X.2021.0080
    Abstract ( 135 )   PDF (7634KB) ( 101 )   Save
    The number of space debris is increasing rapidly which poses severe safety threats to the development of human space activities and in-orbit assets. Among the various proposed active debris removal techniques, tethered spacetug (TST) system has received increasing attention in recent years for its promising application prospects. Some defunct satellites keep specific orientations due to the damage of on-board components and attitude control systems. A dynamics and control study of tethered satellite system during the deorbiting phase was carried out for this special type of large space debris with non-cooperative characteristics. A mathematical model of TST system including two satellites modeled as rigid bodies was developed using Newton laws. According to the different attitude stabilization methods, the emphasis was the spin stabilized satellites and threeaxis stabilized satellites. The control laws of the tug and debris were designed respectively and the influence of the positions of attachment points on debris was also studied. The simulation results indicate that the system will stabilize near an equilibrium state after the countermeasure between the tug and debris.
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    Precise orbit determination of FY-4 synchronous orbit satellite based on optical angle measurement data
    LIU Siyu, HUANG Yong, MAO Yindun, JIA Yaohong, LU Wenqiang, HUANG Chengli, ZHENG Jinghui, YANG Peng
    2021, 41 (6):  54-62.  doi: 10.16708/j.cnki.1000-758X.2021.0081
    Abstract ( 172 )   PDF (5103KB) ( 159 )   Save
    In response to the precise orbit determination and accuracy evaluation requirements of the FY-4 synchronous orbit satellite, the ground optical angle measurement data were firstly used to perform precise orbit determination on the FY-4A satellite. The root mean square of the azimuth and altitude residuals are 0.25″ and 0.45″ respectively. Compared with the orbit determination based only on ranging data, the position accuracy is less than 50m in the arc with angle measurement data. The FY4A satellite orbit was further jointly determined with combined angle measurement data and ranging data, and the orbit overlap accuracy was improved significantly to be better than 15m. Compared with the results of joint orbit determination, the accuracy of real-time orbit products based on only ranging data was evaluated. It is shown obviously that, with the accumulation of ranging data, the orbit accuracy can be gradually improved, e.g., the precision of orbit determination with ranging data can be gradually stabilized to about 20m in 6h.
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    Behavior identification of space adjacent targets
    WANG Yangyang, LIN Bin, YANG Xia, ZHANG Xiaohu
    2021, 41 (6):  63-71.  doi: 10.16708/j.cnki.1000-758X.2021.0082
    Abstract ( 155 )   PDF (5200KB) ( 239 )   Save
    Aiming at the observation behavior of reconnaissance satellite approaching the target by changing orbit, a method of behavior identification for near target based on visual characteristics was proposed. By tracking the feature points of the adjacent target in the image sequence, the trajectory change was obtained. Afterwards, the representation of the possible abnormal behavior of the target in the image was analyzed, and the characteristic parameters were constructed based on the obtained information to judge the target behavior. The experimental results show that the algorithm can process the continuous images under different illumination conditions, and realize the effective identification of the space reconnaissance behaviors with attitude changes of adjacent targets.
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    Parameter identification method of thruster in orbit for deep space probe
    XIN Sibo, ZHAO Xunyou, ZHENG Yiyu, LI Lyuping
    2021, 41 (6):  72-78.  doi: 10.16708/j.cnki.1000-758X.2021.0083
    Abstract ( 143 )   PDF (2703KB) ( 226 )   Save
    According to the needs of China's first autonomous Mars exploration mission, combined with the mass characteristics of the orbiter and the layout of the propulsion system, the influence of jet unloading on the angular momentum of the detector was analyzed. On the basis of the analysis, the angular momentum variation of the detector was calculated by the change law of the rotating speed of three axes before and after the flywheel unloading, and the impulse generated during each jet was solved, as well as the thrust direction deviation. Coordinates of the detector's center of mass were solved through disturbance of each axis from the thrusters within the same group. The jet unloading of China's Mars probe Tianwen-1 in its cruise phase using different thrusters for six times was analyzed with in-orbit data. Calculation results of the thrust vector deviation and the centroid coordinates were highly consistent with design values. The measured thrust vector deviation was not more than 06, and the absolute deviation of the center of mass was less than 18mm, proving that the calculation method was effective and correct which can be used to set the direction of ignition and propellant budget in subsequent orbital control tasks.
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    Health assessment technology of lithium-ion battery for spacecraft based on multi-feature fusion
    YANG Tongzhi, DANG Jiancheng, ZHONG Liang, LIU Yang, LIU Tingyu
    2021, 41 (6):  79-84.  doi: 10.16708/j.cnki.1000-758X.2021.0084
    Abstract ( 153 )   PDF (1882KB) ( 205 )   Save
    The battery is charged and discharged according to the maximum discharge depth in traditional spacecraft battery reliability test. The failure model is used to support the overall reliability design of spacecraft, and cannot be used for the onorbit battery health assessment task. The sample rate, sampling precision and sample size of onorbit spacecraft telemetry are not comparable with those in civil field. The civil battery health assessment method based on high-speed sampling and large sample is not suitable for on-orbit battery health assessment. In order to solve this problem, based on the data characteristics of the on-orbit battery, the degradation features that can be extracted in the on-orbit state were excavated, and a multifeature comprehensive evaluation method was designed to realize the quantitative assessment of on-orbit battery health based on the fusion of degradation features including battery internal resistance, terminal voltage at the same depth of discharge and constant voltage charging time, which can be used as a reference for spacecraft health assessment technology.
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    TopPixelLoss: a loss function for semantic segmentation of remote sensing images with class imbalance
    YUAN Wei, XU Wenbo, ZHOU Tian
    2021, 41 (6):  85-90.  doi: 10.16708/j.cnki.1000-758X.2021.0085
    Abstract ( 222 )   PDF (2754KB) ( 384 )   Save
    Aiming at the problem that the segmentation effect of small target in remote sensing image is not ideal, a loss function named TopPixelLoss was proposed.Firstly, the cross entropy of each pixel was calculated, and then the cross entropy of all pixels was sorted from large to small. After that, a K value was determined. According to the threshold K, the pixels with the largest cross entropy of the top K were selected. Finally, the cross entropy of the K pixels was averaged as the final loss value. Experiments using PSPNet network with cross entropy, FocalLoss and TopPixelLoss were carried out  respectively through Vaihingen data set of ISPRS. The results show that, for different K values, the mean intersection over union (M IOU), F1-score and accuracy(ACC) are all higher than FocalLoss, and that the effect is the best when K is 50000 (MIoU, F1-score and ACC are improved by 3.0%, 5.0% and 0.1% respectively compared with FocalLoss). The proposed TopPixelLoss  function is a very effective loss function for imbalanced class segmentation.
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    Flight program design and implementation of unmanned automatic lunar sampling and return mission
    SHENG Ruiqing, ZHAO Yang, ZOU Leyang, CHEN Chunliang, ZHU Shunjie, HUANG Hao, DU Ying, PENG Jing
    2021, 41 (6):  91-102.  doi: 10.16708/j.cnki.1000-758X.2021.0086
    Abstract ( 158 )   PDF (6698KB) ( 206 )   Save
    In view of the design characteristics and difficulties of Chang'e-5 flight program, such as multi-cabin, multi-target and complex system control, the traditional flight program design method is difficult to meet the mission requirements due to its heavy workload and difficult state control. In this paper, a new method based on state transition was proposed for system modeling of flight program. The whole flight process was firstly decomposed into several module state machines, then each module state machine was divided into state generator, state assessor, command executor and state verifier according to their functions, and modeling design was carried out respectively. Finally, each function module was connected by state generator and state verifier to form the finite state machine description of the whole flight program. Compared with traditional methods, this method has the characteristics of generality, extensibility, reusability, and has great advantages in standardizing flight program design and describing complex flight mission process. The flight program of Chang'e-5 was modeled and designed by using this method, and the design results of typical flight process were given. The in-orbit flight test results show that this method can meet the requirements of the mission and ensure the successful in-orbit flight control of Chang'e-5 mission.
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    Ignition point determination for powered descent phase and in-orbit verification of lunar unmanned sampling and return misson
    ZHAO Yang, SHENG Ruiqing, CHEN Chunliang, ZHANG Xiaowen, ZOU Leyang, GAO Shan, HUANG Hao
    2021, 41 (6):  103-113.  doi: 10.16708/j.cnki.1000-758X.2021.0087
    Abstract ( 116 )   PDF (9170KB) ( 125 )   Save
    Ignition point determination (IPD) is an important part for lunar soft landing process, which is a complex iterative process dealing with orbit design, guidance law design, targeted sampling site determination, landing and lifting-off safety analysis, etc.. IPD may directly influence the final position of landing site and landing safety, and also indirectly influence the realization for lunar surface sampling. In view of the mission requirement of IPD for lunar soft landing, a IPD method by coupling control strategy based onorbit control for multiple times and optimal search on nominal flight track was proposed. Firstly, the principle and constraint condition of IPD were designed and summarized according to unmanned automatic lunar sampling and return mission. Then, the method of IPD for lunar soft landing mission was introduced in detail. Finally, the influence on IPD by major orbit control mission before powered descent ignition of lunar soft landing was analyzed. Meanwhile,the analysis on topographic features and the influence on landing and lifting-off safety of targeted landing region were implemented and then the ultimate ignition point was determined and verified according to the final in-orbit flight results. With in-orbit flight verification of Chang’e-5 mission, the IPD method based on successive approximation and optimization strategy (SAOS) proposed in this paper was effective for this mission and also available for soft-landing mission on celestial body in the coming future.
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    Design and result analysis of the Sun-Earth L1 point exploration
    ZOU Leyang, GAO Shan, ZHAO Chen, QIAO Dezhi, LI Xiaoguang, MENG Zhanfeng
    2021, 41 (6):  114-122.  doi: 10.16708/j.cnki.1000-758X.2021.0088
    Abstract ( 217 )   PDF (4145KB) ( 200 )   Save
    The Sun-Earth L1 point is an optimal location for the Sun observation and is significant for the future Sun observation missions of China. Therefore, the first domestic exploration mission on the Sun-Earth L1 point was designed and implemented during the extended phase of the Chang’E-5 mission. By this flight, the design of the orbit transferring and orbiting around the Sun-Earth L1 point was verified, and the flight environments of the Sun-Earth L1 point such as the TT&C environment, the solar irradiance environment, the three-body dynamics environment and the space radiation environment were detected and verified. The results of orbit flight and environment detection were consistent with the predicted results of the design model and therefore the correctness of the design model was verified with onorbit flight data. Technological achievements, which were important references for the design of deep-space exploration missions and devices in the future, were acquired as expected and further enriched the outcome of the Chang’E-5 mission.
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    Thruster reliability experiment under high temperature on lunar surface
    YU Hangjian, PENG Jing, SHU Yan, ZHAO Jing, JIANG Fan
    2021, 41 (6):  123-131.  doi: 10.16708/j.cnki.1000-758X.2021.0089
    Abstract ( 124 )   PDF (4234KB) ( 87 )   Save
    The landing-ascending assembly of CE-5 experienced strong infrared radiation environment after landing on the moon. Because of the influence of thruster nozzle thermal conductivity, the pipe and solenoid valves faced high temperature environment. The propulsion system suffered oxidant gasification in the pipeline and thruster performance degradation or even inability to work, affecting the normal ascender attitude control after rising from the moon surface. In order to study the performance of thruster at high temperature, the compatibility of fluorine plastic spool and oxidizer at high temperature was investigated, reliability growth and reliability verification test methods for solenoid valves and thrusters were presented, the relationship between temperature and solenoid valve spool stroke was analyzed, the steady state and pulse performance of the thruster at high temperature environment was acquired, and the confidence lower limit of thruster reliability and the reliability of the propulsion system were evaluated under the condition of small sample. The results were verified by high temperature exhaust and the ascent attitude control. The lunar ascent mission was successfully completed. The research methods and ideas proposed have good engineering applicability.
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    Design and implementation of thermal balance test scheme for Chang'e-5 probe
    NING Xianwen, JIANG Fan, CHEN Yang, ZHANG Dong, WANG Yuying, XUE Shuyan
    2021, 41 (6):  132-137.  doi: 10.16708/j.cnki.1000-758X.2021.0090
    Abstract ( 230 )   PDF (4036KB) ( 250 )   Save
    Aiming at the problems faced in the thermal balance test of Chang’e-5 probe, a set of thermal balance test scheme was constructed, and a thermal balance test method was proposed based on the analysis of the current situation of thermal balance test technology of spacecraft at home and abroad, as well as the principle of full verification, effectiveness and comprehensiveness. A special infrared absorption space heat flux simulation method was used. The typical test conditions were designed, and the test technical process was optimized. The results of thermal balance test combined with on-orbit flight show that the thermal balance test scheme can effectively verify the correctness of thermal control design, that the temperature impact caused by the deviation of the space heat flux simulation device does not exceed 2℃, and that the setting of test conditions and the technical process is reasonable. The correlation of thermal analysis model can make the thermal analytical model more accurate and reliable.
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