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We design and complete an experiment of debris detection under the Milky Way background. The data analysis shows that the strong Milky Way background reduces the detector's ability to detect the weak debris, but has little effect on the accuracy of the astronomical positioning of debris. This paper verifies the rationality of using the limit detection magnitude method based on the detection probability to characterize the limit detection ability of detectors. It also verifies the feasibility of using the simulation method to increase the data completeness. The work in this paper will provide a basis for the design of the spacebased detectors and the development of the detection experiments. [ABSTRACT FROM AUTHOR]

The analytic expression of the electrostatic potential near the spacecraft surface is obtained by solving the Poisson equation based on the electrostatic field theory, and the mathematical expression of the electric field intensity near the spacecraft surface is deduced. Using the boundary condition of the electromagnetic field, the dielectric surface potential-permittivity model is established, and the relationship among the dielectric surface potential, surface charge and permittivity is gained. Taking the practical cases for example, the change regulation of the surface charge with surface potential and the distribution of the electrostatic field near the spacecraft surface is discussed, and the effect of the dielectric constant on the surface charge and surface potential is analyzed. The results show that the electrostatic field induced by the dielectric surface charge decreases as the distance incrases when the surface potential of media is constant, and the surface potential of media decreases with the increase of the dielectric permittivity and tends to a stable value under the condition that the dielectric material has equal charge. Therefore, it is possible to select the dielectric materials with smaller permittivity under certain conditions to reduce their surface potential, and to reduce the probability of the electrostatic discharge between different media. [ABSTRACT FROM AUTHOR]

A method for calculating ihe translucent shadow is proposed in this paper for the needs of the design and simulation of the solar array of spacecraft with a large deployable antenna, which can cause translucent shadow. The proposed method builds 3D-model of the metal mesh with the minimum repeating unit to obtain the light transmission rate of the metal mesh under different illumination and builds up a table. When calculating the shadow caused by the metal mesh components, the position relation of the triangular face and the incident light is calculated firstly. The light transmittance of the triangular face is calculated by the table lookup and the light intensity of the projection area is reduced correspondingly. An example of the radial rib antenna shows that the method is fast and accurate. It can generate the light transmission mask pattern correctly, which lays the foundation for the simulation analysis oi the solar array. [ABSTRACT FROM AUTHOR]

Based on the Monte Carlo technology in statistics,combined with the theory of the intersection of a ray with a triangle,the Monte Carlo method for the calculation of Projected Cross-Sectional Areas( PCSA) of spacecraft has been proposed with elaborated procedures. The developed method is parallelized employing a shared memory programming model,Open MP,in order to take advantage of the multiple-core processors,which has been an industry trend for many years. Then,the PCSAs of a one-side circular plate and the SAMSON satellite in some typical attitudes are calculated,aiming at validating the Monte Carlo method and demonstrating its capability of dealing with complex geometries of engineering reguirements. The proposed Monte Carlo method could be applied to the SAMSON and GOCE satellites,with the purpose of computing their PCSAs in large quantities of attitudes,in order to provide input parameters for satellite attitude and orbit control. [ABSTRACT FROM AUTHOR]

The problem of the relative position and attitude coupled control for rendezvous and docking is investigated in this paper, which takes the external disturbances into consideration and assumes that the target spacecraft is freely tumbling. An adaptive 6-DOF control scheme is proposed based on the concept of the fixed-time stabilization. Besides, the singularity problem is solved through a switch function which not only makes the settling time estimable, but also the estimated value is independent of the initial values of states. Then the stability proof is accomplished by the development of a Lyapunov function. The numerical results show that this method achieves good performance and robustness. [ABSTRACT FROM AUTHOR]

This paper studies the impact fluxes calculation for the orbiting space objects via the Chan's model which does not invoke either the kinetic theory of gases model or the Poisson distribution function. Four cases are discussed for the space objects on elliptical orbits passing through a thin spherical shell surrounding the spacecraft. The paper also compares the circular orbit assumption with the elliptical assumption. The satellite Iridium 86 is used to estimate the impact fluxes under both assumptions. It shows that the ratio of the impact flux between the elliptical orbits case and the circular case decreases from about 4 times to 1. 4 times as the thickness increases. The statistical number of the objects gets a big fluctuation as the increased thickness in circular case. The elliptical case is 60 times more than the circular case when the shell' s thickness equals to 5 km. [ABSTRACT FROM AUTHOR]

In order to realize automatization in rendezvous and docking mission, a method of long distance autonomous rendezvous for spacecraft is proposed. A maneuver strategy adapted to long distance autonomous rendezvous is designed. The control parameters are obtained by the autonomous maneuver plan algorithm which contains the initial value calculating and accurate solution iterating. The quasi-mean element method combined with the mean atmosphere density method is used to reduce the on-orbit calculation quantity. The simulation result shows that the design variables of this method satisfy the terminal conditions well. [ABSTRACT FROM AUTHOR]

A shielding effectiveness analysis method based on the equivalent admittance model is proposed to overcome the difficulty of dealing with VHF/UHF shielding of spacecraft. The aperture is treated as a waveguide. The horizontal capacitive model and vertical inductive model are derived from the variation method. The parameter matching admittances and width correction method are also researched. The shielding effectiveness math model is researched based on the transmission line theory. It proves the accuracy of the method and usefulness in dealing with the aperture with arbitrary position and length-wavelength ratio, and it can analyze the resonant effect. The method has been used in analyzing the shielding effectiveness of a telecommunication cavity, and the result comparisons have demonstrated the effectiveness of the approach proposed. [ABSTRACT FROM AUTHOR]

In order to solve the problems of temperature prediction concerned with the spacecraft thermal balance test, a particle swarm optimization algorithm of temperature prediction is proposed in this paper. According to the thermal balance temperature model, the objective function to be optimized and the temperature calculation process of thermal balance based on particle swarm optimization is established. Based on the prophase temperature data of a spacecraft thermal balance test, the thermal balance stable temperature and tranient temperature are predicted. By comparison . the error between the predicted data and the measured data are very little and acceptable. The results show that the proposed method can be effectively applied to temperature prediction of spacecraft thermal balance test. [ABSTRACT FROM AUTHOR]

Prediction method and test validation techniques of vibro-acoustic mechanical environment of large complex spacecraft are presented in this paper. The vibro-acoustic modeling method for spacecraft is discussed based on hybrid finite element-statistical energy analysis (FE-SEA) method, then the satellite-level and systematic-level (fairing-spacecraft- instrument bay-adaptor assemblies) hybrid models are established respectively. In the end, the comparisons between prediction results and acoustic test results have demonstrated the effectiveness of the predication method and hybrid model, and the key issues related to the prediction precision are also discussed. [ABSTRACT FROM AUTHOR]

Single gimbal control moment gyros (SGCMG) system cannot afford accurate three axis torque in singular states. A singularity escape method is designed based on the analysis of singularity principle. Firstly, a new dynamic coordinate system is created and switched as the steering reference coordinate system. Secondly, one of the control dimensions is abandoned when the SGCMG system moves into the singular state and the system is considered as a redundant SGCMG control system within a plane. Thirdly, a null motion in the two dimensional space is provided so that the SGCMG can move out of the singular state quickly. Finally，the three dimensional steering law needs to be switched on after the SGCMG system escapes from the singular state. The method's impressive performance on singularity escape is demonstrated by the simulation results. [ABSTRACT FROM AUTHOR]

To extend the operational life of the failed spacecraft, the technology of spacecraft attitude takeover control is used to take over its attitude control systems. The uncertain dynamics of failed spacecraft and the change of thruster configuration will constitute a formidable technical challenge for attitude takeover control. Therefore, a takeover control strategy based on control system reconfiguration is proposed for the failed spacecraft attitude maneuver. Firstly, a command filtering backstepping control method is used to reconstitute the control law and the stability of the system is analyzed using the Lyapunov function. Moreover, the reconfiguration of the thrusters is implemented by reconstituting the configuration matrix of the thrusters. Furthermore, considering the fuel consumption and the control inputs limitation, the thrust forces reallocation is achieved by the dynamic control allocation method based on constrained quadratic program. The method proposed in this paper achieves the takeover control of the failed spacecraft whose fuel has been exhausted or whose actuators are partially failed. Numerical simulations validate the feasibility of the proposed method. [ABSTRACT FROM AUTHOR]

Single event protection is more and more important in the development of spacecraft. This paper firstly analyzes the characteristics of the impact of single event soft errors on the spacecraft, and then establishes a criticality model for evaluating the impact of a single event soft error on spacecraft. Finally, a weakness identification method and the steps of single event protection are presented, which results provide an objective basis for improving the single event protection design. [ABSTRACT FROM AUTHOR]

Under the influence of the space environment, the organic pollutants applied in the spacecraft have the potential to degrade performance of the sensitive components or devices. In order to control the molecular contamination on-orbit, a method for the zoelite molecular sieve due to its adsorption of organic molecular contamination is proposed in this paper. The absorption capacity of Naβ, NaY and 13X in vacuum is studied with contaminant source from outgassing of cables. It is found that 13X is the best adsorbents compared to Naβ and NaY, and its absorption capacity is 14. 03%. Zeolite-coated cordierite molecular absorbers can be used to collect and retain outgassed molecular effluent emanating from spacecraft materials, thus effectively controlling the contamination of the critical sensitive surfaces to enhance reliability. [ABSTRACT FROM AUTHOR]

Attitude tracking problem with unknown friction torques of reaction wheels and external disturbances is addressed. A robust adaptive controller with parameter update law is proposed and analyzed. An indirect (nonregressorbased) approach used to deal with the effects of unknown friction torque and external disturbances is explored for the attitude control law. This is done by not focusing on friction torques and external disturbances, but on their bound, thus allowing the core information on friction torque and external disturbances to be extracted and used for controller design. The stability of the resulting closed-loop system is proved by application of Barbalat ' s Lemma. Numerical simulations are presented to show the effectiveness of the proposed attitude tracking approach. [ABSTRACT FROM AUTHOR]

To solve the problem of acquiring the load force function description in the process of spacecraft pyroshock environment prediction,a ‘source system-near field structures 'integrated modeling and analyzing method based on Hydrocodes is proposed to analyze the shock processes of the spacecraft-launch vehicle separation mechanisms and other release mechanisms triggered by pyrotechnicsin spacecraft. By using this method,the force time history can be computed,so that the decoupled analysis of the shock can be realized. The analysis of a clamp-band separation structure is taken as an example to illustrate the feasibility of the method. Both the integrated( coupled) model and the decoupled model are constructed and computed,and both comparison and analysis of the computational results preliminarily illustrate the feasibility of the method. Further studies including a series of coupled studies and a simplified model study are also implemented to prove the accuracy and the rationality of partially modeling of the method. In conclusion,this method is a feasible way to acquire the exact function of the pyroshock load force in spacecraft,and provides a basis for the prediction of the spacecraft pyroshock environment. [ABSTRACT FROM AUTHOR]

The arrangement of a heat load with certain temperature in the single-phase active thermal control system that has a low-temperature inner loop, a middle-temperature inner loop and an external loop is discussed, and the arrangement principle is proposed. The total system mass is investigated with Lagrange multiplier method based on the mathematical model of system mass and energy balance relationship. The result shows that there exists a critical heat load temperature at which the total system masses are equal. When the heat load temperature is lower than the critical value the system has lower total mass with the heat load arranged in the low-temperature internal loop. On the contrary, the system has lower total mass with the heat load arranged in middle-temperature internal loop. The heat load arrangement depends on its temperature and power. Finally, the arrangement principle is discussed qualitatively based on the arrangement of temperature difference. [ABSTRACT FROM AUTHOR]