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    25 April 2013, Volume 33 Issue 2 Previous Issue    Next Issue
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    Trajectory Design of Gravity Assist in the High-fidelity Dynamic Model
    Yang-Hong-Wei, CHEN Yang, BAO Yin-He-Xi, LI Jun-Feng
    2013, 33 (2):  1-6.  doi: 10.3780/j.issn.1000-758X.2013.02.001
    Abstract ( 1559 )   PDF (338KB) ( 1179 )   Save
    The precision orbit design using planet gravity assist in the high-fidelity model was studied. Orbit design was divided into two steps, preliminary design and iteration in the high-fidelity model. Preliminary design was based on the simplified model. Conic patched method was used to determine the launch window and the impulse induced by gravity assist. Taking the gravity assist as a continuous process in the high-fidelity model, the swing-by hyperbola was transformed into B-plane parameters. The earth-centered departure impulse was taken as design variable and the problem was solved by differential correction. Orbit design for Flora rendezvous in main asteroid belt after Mars gravity assist was presented and the relation of the results between the two models above was compared and analyzed in the illustrative example analysis. The result indicates that the orbit in the high-fidelity model which is based on the simplified model can be reached after few iterations.
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    Comparison Between Spacecraft Design Load of Random Vibration
    ZHANG Yu-Mei, HAN Zeng-Yao, LIU Shao-Kui
    2013, 33 (2):  7-12.  doi: 10.3780/j.issn.1000-758X.2013.02.002
    Abstract ( 1604 )   PDF (352KB) ( 1361 )   Save
    It is an important step for spacecraft design and analysis to convert random vibration load to quasi static design load, which influences mass of spacecraft structure directly. The theory and methods of design load based on the peak value of acceleration response and displacement response were simply introduced. The differences between the two methods were compared by analytics and FEM. The design load based on the peak value of acceleration response is larger than that based on the peak value of displacement response. The design load based on the peak value of displacement response is recommended and will lighten mass of spacecraft structure.
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    Selection of Integration Center for Chang'e-2 Satellite Extended Mission
    CAO Jian-Feng, HU Song-Jie, HUANG Yong, LIU Lei, LIU Yong, TANG Ge-Shi, LI Xie
    2013, 33 (2):  13-18.  doi: 10.3780/j.issn.1000-758X.2013.02.003
    Abstract ( 2154 )   PDF (338KB) ( 759 )   Save
    The selection of integration center is inevitably involved in the orbital spacecraft calculation, so it is for Chang'e 2  (CE-2)  mission. During its extended mission, CE-2 satellite was placed in the Sun-Earth L2 libration orbit near the weak gravitational area. The selection of the integration center was discussed with theoretical analysis. According to the perturbed twobody problem, it is appropriate to choose the Sun as the integration center. While the force model was calculated by using the JPL DE ephemeris without considering the perturbation of small planets, the Earth would appear to be a better choice.
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    Method of On-orbit Calibrating Scale Factor of Accelerometer Based on Lever Arm Effect
    DANG Jian-Jun, LUO Jian-Jun, WAN Yan-Hui
    2013, 33 (2):  19-24.  doi: 10.3780/j.issn.1000-758X.2013.02.004
    Abstract ( 1623 )   PDF (309KB) ( 1116 )   Save
    Aimed at the low precision and high cost of onorbit calibrating accelerometer scale factor, as well as the auxiliary requirement and the difficulty in engineering implementation, a method of onorbit calibrating accelerometer scale factor based on the lever arm effect was presented. As long as the knob mechanism of the inertial assembly rotates uniformly at the given speed, the scale factor of the accelerometer can be calculated according to the output of the gyroscope and the accelerometer in the inertial assembly. The calibration principle was analyzed, the calibration model was built and the separation algorithm of the error coefficient was derived. The effect of this method was verified by the ground test.
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    Indirect Optimization of Low Thrust Escape Trajectories
    HUANG Gao, HAN Chao
    2013, 33 (2):  25-31.  doi: 10.3780/j.issn.1000-758X.2013.02.005
    Abstract ( 2037 )   PDF (435KB) ( 896 )   Save
    At present, most of the research on low thrust escape trajectory optimization were limited to planar escape or unconstrained engine model. Using the indirect method, the unknown parameters of the two-point-boundary-value-problem were bounded in a hyper unit sphere by multiplying a positive unknown parameter to the fuel optimal performance index. Firstly, the two-dimensional short-time fuel-optimal escape trajectory without control constraint was solved by the homotopy method and curve fits technology. Finally, the three-dimensional long-time fuel-optimal escape trajectory with control constraint was solved. This method is validated to be high-accuracy, good-convergence and efficient for low thrust escape trajectory optimization.
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    On-orbit Calibration Technique Based on the Two-step Moment of Inertia Identification of the Combination Spacecraft
    LIU Wei-Xia, XIONG Zhi, YU Feng, YAO Xiao-Song
    2013, 33 (2):  32-40.  doi: 10.3780/j.issn.1000-758X.2013.02.006
    Abstract ( 1775 )   PDF (465KB) ( 1173 )   Save
    On-orbit identification of the moment of inertia parameters is an important prerequisite to achieve high-precision attitude control of the combination which is combined by the active spacecraft and non-cooperative space target.With that combination,a method for on-orbit identification of inertia parameters in two steps was proposed.The first step,making the moment of inertia of X axis as a benchmark normalized the spacecraft moment of inertia matrix in order to get the special combination of inertia ratio matrix.Then the spacecraft combination attitude dynamics model was established.Based on the angular rate measurement information given by a star gyro,the extended Kalman filter method was proposed to identify all the inertia ratio parameters in about 100s and overcome the shortcomings which the simple model cannot completely identify the rotational inertia parameters information.The second step,according to the identification of the inertia ratio matrix in the first step and using the least squares algorithm,the X axis moment of inertia was obtained.The simulation example verifies the effectiveness of the method,and the identification accuracy is about 1%.
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    Precise Orbit Determination for LEO Satellites Using Single-frequency GPS Observations
    GUO Xiang, ZHANG Qiang, ZHAO Qi-Le, GUO Jing
    2013, 33 (2):  41-46.  doi: 10.3780/j.issn.1000-758X.2013.02.007
    Abstract ( 1712 )   PDF (447KB) ( 1355 )   Save
    To meet the need for precise orbit determination  POD  of the LEO satellites equipped with a single-frequency  SF  GPS receiver and deepen the research on POD with SF GPS data, the problem of cycle-slip detection for SF GPS data was solved, and the reduced-dynamic orbits for HY-2A satellite and ZY-3 satellite were determinated with two different approaches using the SF GPS observations. The precision of the derived orbits was assessed by the analysis of the observation residuals, comparison with the precise orbits derived from dual-frequency (DF) GPS data and validation with satellite laser ranging  SLR  measurements. The results indicate that the 3D precision of the SF orbit without ionospheric delay correction is between 2dm and 3dm for HY-2A and about 1m for ZY-3, and the notable improvement with the group and phase ionospheric correction GRAPHIC observation is achieved to eliminate the first order ionospheric delay and the 3D precision can reach 1dm for HY-2A and 1~2dm for ZY-3. The results demonstrate that the sub decimeter level accuracy can be achieved for POD of LEO equipped with a SF GPS receiver.
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    Study of  Two-phase Flow in Injector Tube of Monopropellant Thruster
    CHEN Jun, WANG Meng
    2013, 33 (2):  47-53.  doi: 10.3780/j.issn.1000-758X.2013.02.008
    Abstract ( 2071 )   PDF (420KB) ( 1257 )   Save
    During the application of satellite propulsion systems, the high temperature raised in the process of monopropellant thruster is transferred to the injector tube, leading to the two-phase flow phenomenon. As the massflow of propellant decreases, the pneumatic resistance will increase accordingly. By using the FLUENT software, the influence of pneumatic resistance on the thruster performance was studied. Based on the proportion fraction model and government equations of multiphase flow, the flow and pressure distribution in the injector tube under 84 operating conditions were presented. Furthermore, special tests on multiphase flow-combustion pressure relationship were also performed. Experimental and numerical results show that the pneumatic resistance can result in the sudden variety of pressure drop when the inner diameter of injector tube is small enough, and which cause the decreased thrust ultimately.
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    Optimal Trajectory Design for the Lunar Module in Ascent Stage
    MA Ke-Mao, CHEN Hai-Peng
    2013, 33 (2):  54-60.  doi: 10.3780/j.issn.1000-758X.2013.02.009
    Abstract ( 1637 )   PDF (312KB) ( 797 )   Save
    To realize the trajectory optimization of the lunar module in ascent stage, the dynamical model of the lunar module was established and nondimensionalized to construct the optimal control-oriented model. Based on Pontryagin's minimal principle and taking the fuel consumption as the optimizing index, the optimal problem was converted into a time free two-point boundary value problem ( TPBVP) . Using one of initial value hypothesizing methods, combined with the forward sweep algorithm, the TPBVP was solved and the optimized trajectory was obtained. Numerical simulation was conducted. The simulation results justify the rapid convergence rate and high reliability of the proposed method.
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    Design of Multiple Gravity Assist Trajectories with Deep Space Maneuver
    ZHANG Ying, YUE Xiao-Kui, HE Liang
    2013, 33 (2):  61-66.  doi: 10.3780/j.issn.1000-758X.2013.02.010
    Abstract ( 1463 )   PDF (357KB) ( 1192 )   Save
    Planetary swing-by is an effective way to reduce the launch energy for interplanetary exploration missions. However, the traditional swing-by model is difficult to match the relative velocity vector before and after the swing-by. Therefore, the swing-by model was established, and the hyperbolic velocity vector after the swing-by was also derived. Because of the strong coupling of the trajectory design parameters, a global-local hybrid search algorithm was proposed for designing Earth-Venus-Earth-Mars-Jupiter transfer trajectories. All these results certify the effectivity of the proposed model.
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    Satellite Attitude Determination Based on the Adaptive Federated Kalman Filter
    LI Peng, TANG Jian, DUAN Guang-Ren, SONG Shen-Min
    2013, 33 (2):  67-71.  doi: 10.3780/j.issn.1000-758X.2013.02.011
    Abstract ( 1642 )   PDF (292KB) ( 1234 )   Save
    Standard Kalman filter adopts constant covariance of measurement noise. When statistical characteristics of measurement noise changes, estimation error increases, which results in filtering divergence. An adaptive federated Kalman filter was proposed with fuzzy adaptive Kalman filter but not Kalman filter in the subsystem of federated Kalman filter, and the weighted coefficient of covariance matrix was adjusted by fuzzy inference algorithm real-timely. It made the measurement noise of the dynamic equation close to the truth level. When it is applied to multi-sensor attitude determination systems, simulation results demonstrate the true effectiveness of the proposed algorithm.
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    Momentum Management of Chang'e-2 Satellite on L2 Point
    DAI Ju-Feng, XU Hong-Bing, CUI Yan, XUE Rui
    2013, 33 (2):  72-77.  doi: 10.3780/j.issn.1000-758X.2013.02.012
    Abstract ( 2789 )   PDF (312KB) ( 824 )   Save
    Since Chang'e-2 (CE-2) satellite entered the Lissajous orbit surrounding the EarthSun Lagrange point L2,the effect of jet uninstall was studied. A momentum management method by light pressure was given. The in orbit test shows that the Solar light pressure is strong enough to unload the momentum of CE-2 satellite, which is surrounding the L2 point. This method can substantially reduce the number of jet uninstall, and benefit the orbit maintenance of CE-2 satellite.
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    Research for Lunar Dust Effects and Its Ground Simulation Methods
    TONG Jing-Yu, LI Man, BAI Yu, TIAN Dong-Bo
    2013, 33 (2):  78-83.  doi: 10.3780/j.issn.1000-758X.2013.02.013
    Abstract ( 1438 )   PDF (424KB) ( 1300 )   Save
    The character of lunar dust and its effects on lunar exploration were introduced. Based on analysis of lunar environment and its effects on lunar explorer, such as contamination, abrasion, choke and static, some preliminary propositions for key research topics were given,including lunar dust environment simulation, the causes of lunar dust risen, the kinematics and dynamics characteristic of lunar dust, lunar dust mitigation and test evaluation methods. According to the requirements of these topics, the principal environments for a complete lunar dust simulation facility was generalized, such as vacuum, temperature, solar wind, solar ultraviolet, lunar dust, electric field and magnetic field of lunar surface. Then the project of a new lunar dust ground simulation facility was designed. The test methods for lunar environment effects were discussed preliminarily. The proposed method can give a reference for facility development and test research in future lunar exploration.
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