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    25 February 2011, Volume 31 Issue 1 Previous Issue    Next Issue
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    Semi-empirical Method for Force Specification in Force Limited Vibration Test
    LI Zheng-Ju, MA Xing-Rui, HAN Zeng-Yao
    2011, 31 (1):  1-7.  doi: 10.3780/j.issn.1000-758X.2011.01.001
    Abstract ( 4536 )   PDF (372KB) ( 2569 )   Save
    The force limited vibration test can alleviate the over testing phenomenon generated by the resonant frequency in the acceleration-controlled vibration test. Force specification estimation is a prerequisite for the force limited vibration test. The principle of force limited vibration test was introduced. And the calculation method for semi-empirical coefficient was derived by the two degree of freedom system. Furthermore, the force specification of some spacecraft payload was caculated in the random vibration test. The results show that this method is reasonable to estimate the force specification.
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    Orbit Design for Mars Exploration by the Accurate Dynamic Model
    CHEN Yang, ZHAO Guo-Qiang, BAOYin He-Xi, LI Jun-Feng
    2011, 31 (1):  8-15.  doi: 10.3780/j.issn.1000-758X.2011.01.002
    Abstract ( 4981 )   PDF (479KB) ( 2617 )   Save
    The precision orbit design for Mars exploration by the accurate dynamic model was studied. The launch window and trans-Mars orbit was determined through the partical swarm optimization (PSO) algorithm within the heliocentric two-body restriction. The patched conics method was introduced to design the earth-centred parking orbit and departure hyperbolic orbit. The solution of two-body Lambert problem was input as initial value for precision orbit design, and the preliminary orbit was corrected with the restrictions of the Mars B-plane parameters and flight time by the accurate dynamic model. Finally the designed orbit was simulated with the STK softwares.
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    Decentralized Attitude Coordination Control of Satellites within Formation under Input Constraints
    ZHANG Bao-Qun, SONG Shen-Min, CHEN Xing-Lin
    2011, 31 (1):  16-24.  doi: 10.3780/j.issn.1000-758X.2011.01.003
    Abstract ( 3900 )   PDF (470KB) ( 2338 )   Save
    The coordination control problem for formation flying satellites associated with time-varying reference attitude tracking was investigated, and a nonlinear, saturated, and decentralized coordination controller was proposed. By introducing a hyperbolic tangent function vector in the controller, the boundness of the continuous control input was guaranteed. Barbalat′s lemma was employed to analyze the stability of the closed-loop attitude coordination controlled system, then the result of the asymptotical stability was obtained. Simulations under various conditions validate the effectiveness of the proposed algorithm and establish the relation between the relative-attitude-keeping performance during the transition and the factors, which are the topology of the information communication graph of formation satellites and controller gains.
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    External Heat Flux on Manned Transport Spacecraft with Multiple Modes and Attitudes
    LU Wei, HUANG Jia-Rong, ZHONG Qi
    2011, 31 (1):  25-32.  doi: 10.3780//j.issn.1000-758X.2011.01.004
    Abstract ( 3427 )   PDF (483KB) ( 2110 )   Save
    External heat flux analysis is not only the foundation of thermal control design and thermal analysis, but also the significant thermal boundary condition for ground thermal test. Based on theoretical analysis, a spacecraft external heat flux model was developed and the heat flux was calculated in different flight modes and attitudes. In addition, the heat flux characteristics was obtained in the extreme case. The results show that heat flux increases with the augmentation of percent time in sunlight when the spacecraft is in three-axis stabilized attitude, but decreases abruptly when it turns into the yaw maneuver, and then the heat flux will decrease with the augmentation of percent time in sunlight reversely.
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    Target Spacecraft Phasing Strategy in Orbital Rendezvous
    ZHANG Jin, HUANG Hai-Bing, WANG Wei, TANG Yi
    2011, 31 (1):  33-41.  doi: 10.3780/j.issn.1000-758X.2011.01.005
    Abstract ( 4581 )   PDF (447KB) ( 2394 )   Save
    To make preparations for the rendezvous and docking, the target phasing maneuver mission is a non-fixed time rendezvous problem. According to the basic phasing principle, the phase influences by atmosphere drag, drift of right ascension of ascending node and orbital errors were analyzed and used to improve the approximate formulae, which estimated the effect of impulse on the phase angle. With the combination of the approximate formulae and the numerical integration, the target phasing strategy was improved in three aspects: the calculation of maneuver data for the perturbed trajectory, the selection of the phasing-direction and the selection of the maneuver revolution number. The simulation results show that the proposed target phasing strategy can obtain the maneuver data for the perturbed trajectory and can meet the terminal precision requirements. Compared with the phasing-direction selection method, with which only the terminal phase angle deviation is considered, the proposed strategy can reduce the fuel consumption.
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    Transfer Matrix for Relative Dynamics in Elliptic Orbit
    YUE Xiao-Kui, YUAN Yun-Xia
    2011, 31 (1):  42-47.  doi: 10.3780/j.issn.1000-758X.2011.01.006
    Abstract ( 4536 )   PDF (283KB) ( 2866 )   Save
    Based on Lawden equation, a new form of state transfer matrixes was presented. The state transfer matrix was derived in terms of position-velocity in solid space, and transformed from angle domain to time domain, so the state transfer matrix denoting true position and velocity was obtained. Arbitrary forms can be obtained by primary transformation. The numerical simulations show this state transfer matrix is effective and feasible for theory and project applications. It can be used for solid orbit rendezvous, orbit correction, formation configuration design and control for elliptic orbits with arbitrary eccentricity.
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    Error Analysis for Rendezvous and Docking with Linear Covariance Method
    WANG Da-Peng, LIU Yu-Qiang, CHEN Shao-Long, LIU Yong-Jian
    2011, 31 (1):  48-55.  doi: 10.3780/j.issn.1000-758X.2011.01.007
    Abstract ( 3775 )   PDF (486KB) ( 2408 )   Save
    Based on rendezvous mission and flight characteristics, a rendezvous orbit consists of three basic flight segments, i.e., orbital maneuver, free flight and midcourse velocity correction. The propagation laws of state errors and effects of errors on the accuracy of the rendezvous and docking (R&D) mission were studied with a linear covariance analysis method and Monte Carlo simulations. In the orbital maneuver segment, effects on the state error propagation caused by attitude errors, control system performance, state estimation errors of the chasing spacecraft, and orbit perturbation of the target spacecraft were studied. In the free flight segment, effects on the state error propagation caused by prior state estimation errors and orbit determination errors of the chasing spacecraft were studied. And in the midcourse velocity correction segment, effects on the state error propagation caused by attitude errors and control system performance of the chasing satellite  were studied. The results of the Monte Carlo simulations show that the proposed error analysis  method is properly designed and can be applied to the R&D orbit design and error analysis so as to validate the R&D strategy.
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    Reliability Verification Test Methods of a Connecting/Separating Mechanism for Manned Spacecraft
    LIU Zhi-Quan, SUN Guo-Peng, GONG Ying
    2011, 31 (1):  56-61.  doi: 10.3780/j.issn.1000-758X.2011.01.008
    Abstract ( 3989 )   PDF (406KB) ( 2269 )   Save
    The functions, principles and failure modes of a connecting/separating mechanism for manned spacecraft were introduced. Reliability characteristic parameters of the mechanism were given. The reliability verification test method of the mechanism was put forward, including test method choosing, test state determination, test program, failure criterion and  reliability  assessment  method. An application example of the method was shown. By these reliability  verification  test methods, the reliability of the connecting/separating mechanism for manned spacecraft could be verified.
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    Thermal-structural Modeling and Simulation for Scan Mirror on Imager in Solar Radiation
    YOU Si-Liang, CHEN Gui-Lin, WANG Gan-Quan
    2011, 31 (1):  62-69.  doi: 10.3780/j.issn.1000-758X.2011.01.009
    Abstract ( 3978 )   PDF (646KB) ( 2357 )   Save
    The solar-radiation disturbance exists in the optical payloads onboard the three-axis stabilized satellite platform. It leads to thermal distortion, which obviously affects the imaging quality of optical system. The approximate modeling technique of complex external heat flux was applied, as well as the multi-software-coupled finite element simulation. The thermal effect of the imager′s scan mirror caused by solar radiation was researched, and compared under different thermal design conditions on the FY-4 imager. The optimal scheme of the thermal design of the scan mirror was obtained. The result shows that the thermal distortion is closely related to the thermal coating on the side and the back of the scan mirror. It is difficult to  resolve  the conflict  between the temperature fluctuation and the thermal distortion if only using the thermal design method. It is necessary to improve the material of the scan mirror or adjust the optical-mechanical structure.
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    Integral Filter Method of Orbit Determination for Geostationary Satellite Based on Spaceborne GNSS
    LIU Li, DONG Xu-Rong, ZHENG Kun, WANG Wei
    2011, 31 (1):  70-75.  doi: 10.3780/j.issn.1000.758X.2011.01.010
    Abstract ( 3765 )   PDF (315KB) ( 2558 )   Save
    Spaceborne global navigation satellite system (GNSS) method is used to overcome the problem of poor satellite visibility  when using GPS for orbit determination of geostationary orbit (GEO) satellite. For FENGYUN geostationary satellite, the visibility of GEO satellite relative to GNSS in the future was analyzed. Since the sampling interval of navigation receiver is a little  longer, an integral filter method, combined with the orbit integral and Kalman filter methods, was proposed to determine the geostationary satellite orbit. After generating the necessary data by STK software and programming by MATLAB software, a simulation test was made. The simulation results  show that the performance of the proposed method is good and that accuracy of orbit  determination  for GEO satellite is higher.
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    Rotating Formation Flying for Tethered Micro-satellites
    CHEN Zhi-Ming, LIU Hai-Ying, WANG Hui-Nan, FENG Cheng-Tao
    2011, 31 (1):  76-83.  doi: 10.3780/j.issn.1000-758X.2011.01.011
    Abstract ( 4541 )   PDF (451KB) ( 2326 )   Save
    A tethered control method was studied in order to solve three satellites formation  flying problem. Based on the equilibrium analysis of spinning rigid body satellite, the Thomson and Likins-Pringle equilibrium modeling of three satellites formation flying using tethers were established. The equilibrium conditions were obtained through equilibrium analysis of formation system, and then three control methods were presented to solve disequilibrium problem of Likins-Pringle equilibrium system. Simulation  results demonstrate that the Thomson equilibrium system will be in steady-spin motion with specific conditions, and Likins-Pringle equilibrium system will operate steadily using spring system and tethers added with thrusters control methods under specific conditions.
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