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To enable the plug-and-play of the spacecraft systems, a middleware software system should be equipped beforehand to automatically deal with the self-discovering, self-configuration and communication between the components. A SPA software solution based on the Robot Operating System (ROS), or RSPA in short, was proposed. The RSPA software architecture that can handle the heterogeneous nature of SPA network was presented after the analysis and comparison between the software architectures of ROS and SPA. Then the problem of the implementation of SPA protocol through ROS environment was addressed. Finally, an emulation system which was used to verify the effectiveness of RSPA was demonstrated. Compared with the previous space plug-and-play architecture software implementation, RSPA has a clear network model, is easier to expand, and can effectively utilize ROS′s stable infrastructure and rich tools to accelerate software development. [ABSTRACT FROM AUTHOR]

Each module of China Space Station has its own power system. Due to the sunshade effect, some modules can be in the state of power shortage. To solve this problem, the scheme of using electrically isolated high power supplement controller ( PSC) is proposed. The PSC is modularly designed with basic converter unit (CU),the output capacity of PSC can be increased by CU parallel. This paper mainly describes the power bus electrically isolated conversion topology, double-loop of constant voltage and constant current control, maximum current sharing control, output digital regulation, and the design of Fault detection, Isolation and Recovery (FDIR). On this basis, PSC realizes high reliability power supply between different modules of China Space Station with the 4kW power level, and the scheme is proved by the experimental results. [ABSTRACT FROM AUTHOR]

In order to improve the payload ratio of spacecraft and reduce the weight of the platform .this paper presents a new way to reduce the harness weight by using matrix control and telemeter harness. To optimize the harness design method to the greatest extent, a mathematical model of matrix harness based on undirected graph was established.Two factors affecting harness weight lose were extracted by the model： topology and matrix resource distribution. The concept of visiting order was introduced in the paper， and the organic combination of the two factors was realized. Basic principles for optimization design were presented,and simulation verification was carried out.The results show that the optimized matrix harness weight is about 50% less than the discrete harness weight， improving the system reliability effectively. [ABSTRACT FROM AUTHOR]

High-precision and all-time spacecraft autonomous navigation is the development direction in the space technology application, and is also the foundation for the actual onorbit application for the spacecraft. An autonomous navigation algorithm for spacecraft based on the astronomical information of the sun, the earth and the moon was researched. Firstly, by using the dynamics equations and the angles between the spacecraft, the earth, the sun and the moon, as well as the distance between the spacecraft and the earth as the state model and observation model respectively, the navigation system was established. Then, the autonomous navigation algorithm was presented. When the sun and the moon were observable, the autonomous navigation through the nonlinear cubature Kalman filter (CKF) was adopted: and the high-precise orbit prediction algorithm was used to predict the orbit by using the track dynamics directly during the eclipse. Finally, the numerical simulation was provided. Results show that the positioning accuracy of this method is lower than 2km. which is enough to satisfy the autonomous navigation. [ABSTRACT FROM AUTHOR]

Based on the analytical solutions of T-H equations and its state transition matrix form, the open-loop control method of spacecraft impulsive relative hovering was studied， which is promising for practical engineering use. The true anomaly intervals of the hovering impulse were optimized by the nonlinear mathematical programming. Based on the calculation of collision probability, the method of safety analysis and risk management was proposed. The numerical simulations show that the introduced relative hovering method can be used for circular and elliptical reference orbits hovering. Furthermore, the local optimal solution can be obtained by applying the true anomaly intervals optimization method. The maximum collision probability and the minimum relative distance nearly appear at the same time. And, the smaller the relative distance is, the larger the collision probability. [ABSTRACT FROM AUTHOR]

The problems of momentum envelope analysis and momentum management for spacecraft with a redundant flywheel array were investigated. Firstly, based on analysis of the visualized envelope, a characteristic parameter set was established, and an algorithm was proposed to determine the maximum momenta available from the wheels in a specified direction in three-dimensional total momentum space. Furthermore, a momentum management strategy was developed based on a null motion allocation taking the maximum momenta as the reference, to fully utilize the momentum envelope capacity and to avoid the case where wheel momentum saturation appears before the total momentum arriving at the envelope. Finally, comparative numerical simulations were provided to illustrate the effectiveness of the proposed momentum management strategy. [ABSTRACT FROM AUTHOR]

Based on a state transition matrix, a method for spacecraft multiple-pulse hovering was presented. The relative motion dynamics model and the state transition matrix were introduced to describe the relative motion of two spacecraft in circular, elliptical and hyperbolic reference orbits. A method of achieving the orbit-transfer velocity increment for hovering was deduced using the state transition matrix. The numerical simulation shows that the proposed method can achieve spacecraft relative hovering effectively. With the increasing of impulse number, the precision of relative hovering is promoted. Moreover, the precision of relative hovering is influenced by the eccentricity of reference orbit. [ABSTRACT FROM AUTHOR]

There is limitation to confirming the reason for the data change of attitude sensor of spacecraft after the telemetry data were received, which brings the problem that the data can′t be analyzed in time. To solve this problem, an approach was proposed to study the range of solar disturbance for attitude sensor based on two angle parameters which reflect relative position between the sun and spacecraft. The range of solar disturbance for attitude sensor could gain in advance due to well-regulated change of the two parameters. Consequently, which could offer help for data change and failure analysis of attitude sensor. Furthermore, it is beneficial to the design of spacecraft attitude control system, for example, configuration and installation of attitude sensor. In addition, the faculty to restrain sunlight of attitude sensor is validated through the way. Simulation results prove the validity of the algorithm. [ABSTRACT FROM AUTHOR]

computation model integrated with interrupts and tasks was presented for the analysis of worst case response time (WCRT) for an aerospace embedded system with multilevel interrupts. Both interrupts and tasks was used for a unified priority definition in this model, so the timing analysis of nested interrupts can be combined with task response time analysis techniques. The WCRT of both interrupts and tasks was derived in the integrated model. Furthermore, detailed descriptions of key parameters were proposed and the differences between interrupts and tasks were also discussed. The open source LE0N3 processor platform and Modelsim software were introduced to simulate and verify the method. The results show that the overestimate of task WCRT is less than 5% which demonstrates the effectiveness of the method and the value for engineering applications. [ABSTRACT FROM AUTHOR]

In order to improve the temperature control precision of the detector and other spacecraft and reduce the temperature fluctuation accuracy, a one-way micro expansion type heat switch was designed based on thermal expansion and contraction of metal. The characterization parameter was get to analysis the two-dimensional steady-state thermal modeling. Then, an experimental setup was designed to simulate the actual working conditions of thermal .switch, and switch operating status and related performance parameters were measured under different heat power. When the power is up to 0.4W. disconnected resistance is 230.46 K/W. The power is up to SOW, the closed thermal resistance of switch is 0.302 2 K/W, and switch ratio is 762. 6. The results shown the experiment data are in good agreement with the theoretical calculations. [ABSTRACT FROM AUTHOR]

To improve the temperature control reliability of spacecraft thermal control system, a self-adaption temperature control cold plate was designed based on the volume change characteristics of the paraffin wax due to the phase-change as its temperature changed. The design model of the self-adaption temperature control cold plate was built, and the thermal net model of the single phase fluid loop with the cool plate was given. The performance of the cold plate and the single phase fluid loop was analyzed. Results show that the cold plate designed has the ability of self-adaption temperature control, and it may be widely used in the temperature control for the electronic devices at ground and heat loads on spacecraft. [ABSTRACT FROM AUTHOR]

The problems of approach planning and control method for rendezvous with an uncontrolled tumbling non-cooperative target in an elliptical orbit were studied. On the part of approach planning, the chaser and target spacecraft rendezvoused from a distance to a safe berth point, which was designed by considering the structure of the target spacecraft. After the observation at the berth point for a period of time, the grasp point was determined and the attitude motion of the target spacecraft can be predicted. Then, a safe approach trajectory from the berth point to the implement point for grasping was designed for tracking. On the part of control method, by using the measurements of the relative position and considering the uncertainties of the actual system, adaptive control laws based on the characteristic model were proposed for the whole rendezvous and approach process. Finally, an example was given to show the effectiveness of the proposed approach planning and control method. [ABSTRACT FROM AUTHOR]

For investigating the effects of passive vibration isolator parameters on the vibration transmissibility, nonlinear stiffness characteristics of the vibration isolating material was described by the cubic polynomial function of its deformation, and the nonlinear damping was characterized by viscoelastic fractional derivative operator. The nonlinear dynamic model in microgravity environment of passive vibration isolator system was developed. The dynamic response characteristics were analyzed by the harmonic balance method, and the vibration transmissibility was obtained. Then the influence of stiffness? damping and mass of the isolator and isolated body on vibration isolation effect were analyzed. The results provide theoretic reference for design of spacecraft isolators. The results show that the effects of vibration isolator nonlinear damping coefficient on vibration transmissibility were obvious and the effects of vibration isolation object mass on resonance vibration peak value have a great deal to do with the vibration isolator nonlinear damping coefficient. [ABSTRACT FROM AUTHOR]

The free-floating manipulator system with non-zero initial momentum was studied. The configuration parameters of the manipulator system and the Lagrange equations were used to build the dynamic model. In order to eliminate the coupling effects between the base and the manipulators in angular velocity, the zero reaction motion equation was derived by using the manipulator system redundancy. A polynomial interpolation method was used to approach the trajectory of end-effector and the trajectory was planned due to the initial and final position of end-effector which could turn workspace of end-effector into joints space. A suitable proportion- differentiation (PD) control algorithm was designed to drive manipulators movement. Simulation results demonstrate that the end-effector can smoothly track the target and the movement of the manipulator doesn't make any disturbance to the base's attitude at the same time and the spacecraft also can work normally because of the stability of base. [ABSTRACT FROM AUTHOR]

Accurate measurement of the relative position and attitude between the active spacecraft and the target spacecraft is the key to the successful completion of rendezvous and docking (RVD) mission. In the traditional vector measurement methods, the position and attitude are usually treated separately, which led to the difficulty of algorithm design and computational complexity. A new binocular vision measurement algorithm based on motor algebra was proposed to uniformly describe the relative position and the attitude of the two RVD spacecraft. This new algorithm has a more concise form and some computational advantages compared to the traditional methods. The relative position and attitude determination problem was simplified to solving two linear equations. Simulation results show that the algorithm can meet the measurement accuracy requirements of RVD missions. [ABSTRACT FROM AUTHOR]

A modeling method of reverberant load which meets the spatial correlation was proposed for the vibro-acoustic analysis of spacecraft, which can simplify the modeling process of vibro-acoustic analysis and be convenient for engineering applications. Firstly, the spatial correlation theoretical model of reverberant load was established based on wave theory, then combined the requirements of spacecraft engineering, the reverberant load model was developed on the basis of Rayleigh integral and reciprocity relationship, and the satisfaction with spatial correlation theoretical model was proved. Finally, the availability of this method was guaranteed by typical system analysis, and the effect of spatial correlation of reverberant load on the vibro-acoustic analysis was evaluated. [ABSTRACT FROM AUTHOR]

In order to further improve the reusability of the on-board software and to solve the problem of the on-board software inefficient development, the characteristics of the spacecraft software and the software component technology were analyzed. Then a definition of the on-board software component was given, and an on-board software component model suitable for spacecraft was proposed based on the non-object oriented language. The on-board software components are independent, integrate, assemble and functional. Besides, the method for the on-board software component design, develop and implement was described. It can be the guide of the on-board software component development based on the non-object oriented language. Some examples were given, in which the software systems were developed according the proposed method. The results show that spacecraft software component development methods are effective, and the spacecraft software development efficiency can be increased by 200% ~ 500%. [ABSTRACT FROM AUTHOR]

High temperature heat pipe receivers are commonly used as a core component of space solar dynamic power system. On the base of heat conduction governing differential equation under microgravity, two-dimensional numerical analysis of PCM container was conducted by a enthalpy method. With both void cavity and phase change considered, thermal performance of heat pipe receiver was analyzed. Numerical results were compared with NASA results. The accuracy of calculation model under gravity was verified. The results indicate that void cavity influences the process of phase change. The void cavity reduces the utility and thermal storage ability of PCM. The temperature gradient of PCM zone is very significant because of the void cavity. The PCM contained in the integrated heat pipe performs the averaging function of the heat loads. Normal working of wick ensures the uniformity of heat pipe, thus heat pipe receiver alleviates thermal spot and thermal ratcheting. [ABSTRACT FROM AUTHOR]

A reliability validation test method and a conservative reliability assessment method for pyrotechnic cutters were put forward based on characteristic parameter measurements with small samples. As characteristic parameters, the peak gas-pressures of the cutter in the rated condition (bilateral ignition) and in the unilateral ignition condition were measured in the test. Based on the generalized strength-stress theory, the reliability assessment was carried out by test data with small samples. The proposed reliability validation test method not only avoids consuming more test products to obtain critical value distribution of the characteristic parameters, but also utilizes characteristic parameter variables instead of counting the success number of samples. Besides, This method doesn' t need artificial intensification coefficients. Hence, the reliability assessment results can reflect the real reliability level of the products in the worse working conditions. An application example was given, in which only 60 test samples of the pyrotechnic cutter were used for the reliability validation test and the reliability assessment. The method contributes to the small sample reliability assessment of pyrotechnic devices in spacecraft. [ABSTRACT FROM AUTHOR]

Currently, the most commonly used pattern in space target observation of space-based optical system is natural intersection mode which needs a long coverage period, and the effect of atmosphere radiation of the earth on space target observation is not estimated. Based on the characteristics of space-based platforms on sun-synchronous orbit, several observation modes for GEO targets were planned and analyzed, transforming the effect of atmosphere radiation of earth on space target observation into a camera off-axis angle constraints, which is valuable for engineering application. The observation efficiency was evaluated from several aspects such as attitude modes, coverage time and observing arc. Both theoretical analysis and simulation results show that the regional staring mode has a big advantage for GEO object observation. The coverage time is less than 24 hours, while the attitude mode is comparatively simple, and the observing arc is longer. [ABSTRACT FROM AUTHOR]

The noise will directly affect the accuracy of the ultrasonic flow meter. The sources of noise inside the small pipe ultrasonic flow meter were analyzed. Considering the critical real-time measurement for the flow meter, a method of directly denoising the transit-time difference data was proposed. It avoided sampling and analyzing the original ultrasonic signal, which might lead to the poor performance of real-time measurement. For different noise sources, two different kinds of filters were applied. The wavelet filter was chosen as the first denoisng filter, which used a hard thresholding to reduce error from the random noise. A median filter and a mean filter were combined as the second denoisng filter to reduce error from the acoustic noise. The flow rate measurement experiment was implemented to verify the validity of the proposed denoising method. The results show that the instantaneous measurement error can be less than 1% after denoising. [ABSTRACT FROM AUTHOR]

The high gainantenna was drived with the stepper motor, the harmonic gear and the microstep driver. The over all antenna pointing error shall be 0.1° (3σ) or less. The model of driving antenna was described. On account of the torque model of stepper motor, the nonlinear dynamics model of the harmonic gear and the dynamic model of the microstep driver, as well as the nonlinear model was built, the step response, pointing error and output torque were provided by the proposed model. Comparing with the simulation and experiment results, the math model is correct and reliable. The pointing error is 0.017° (3σ) or less,m eeting the requirements for pointing precision. [ABSTRACT FROM AUTHOR]

To maintain the three-axis attitude of the underactuated nadir-pointing spacecraft, an attitude hold control method based on the pitch-axis bias momentum wheel and roll-axis thruster was proposed. The method was based on coupling between the roll and yaw axes of biased-momentum spacecraft, and could avoid the problem that zero momentum spacecraft s underactuated axis and actuated axis were almost decoupled. Spacecraft attitude dynamics was divided into long-period dynamics and short-period dynamics, and the control laws were designed by the pole-assignment method. The scale of control parameters was confirmed, and the stable state error of the control system was obtained. The simulation results show that the designed controller can eliminate initial attitude error of the spacecraft and resist the external disturbance, and the spacecraft attitude is maintained at the equilibrium point. [ABSTRACT FROM AUTHOR]

Aiming at the typical problems of electromagnetic pulse influencing wires of the spacecraft, the spacecraft field/wire coupling analysis model was built. To calculate the electromagnetic field, the finite difference time domain method was improved. The average electrical parameters was adopted, and the object edges was modified, which cannot match the Yee grids. To analyze the field/wire coupling, the multiconductor transmission line transient reponse model based on the lumped equivalent sources was proposed. The distributed voltage and current sources excited by the incident field were decoupled with the exponential multiconductor transmission lines matrix. The lumped equivalent voltage and current sources model were built. The model was used to analyze a wire in the launch vehicle which was influenced by the electromagnetic pulse. The analysis results show that the interference of the electromagnetic pulse to the wires can destroy the equipment, even though there is shield. [ABSTRACT FROM AUTHOR]

To study the influence of the void on phase change heat transfer process under microgravity, the void model using temperature sorting algorithm was built based on the enthalpy method, which can simulate the interaction between the void and the temperature distribution. The 2D model of phase change material container was built to study the void migration rule under the periodical heat flux boundary condition. The results show that from the initial time to the 8th period the void disappears gradually in the low temperature region and appears in the high temperature region. The void moves to the high temperature region and diffuses along the isotherms. The void migration increases heat resistance along the heat transfer route and the temperature difference about 3 °C between cold and hot boundaries. [ABSTRACT FROM AUTHOR]

The Unified S-band System (USB) is the main TT&C (Telemetry, Tracking and Command) network used in space missions including manned spacecraft and lunar exploration programs of China. Firstly, all kinds of errors existed in the USB range and range-rate measurements were analyzed, and the error correction models of tropospheric refraction delay, ionospheric delay, time tag bias, phase center error, channel delay and satellite ephemeris error were presented. Secondly, the accuracy of the USB measurements was evaluated and obtained based on the residual analysis of the USB range and range-rate measurements of a real space mission. Finally, according to the current actuality of disposal and application of USB measurements, some advices that should be noted in data processing were put forward. The conclusions have important reference value to improve the accuracy of orbit determination based on the USB system of China. [ABSTRACT FROM AUTHOR]

The main objective of spacecraft TT&C (Track, Telemetry & Control) resources scheduling problem is to maximize the spacecraft tasks through TT&C resources allocation with some scheduling methods. A time window-oriented model was proposed in reference to the task-oriented models. Furthermore, the branch and bound algorithm of the model and its complexity analysis were given. Finally, the validity and practicability of the algorithms were showed by simulation. [ABSTRACT FROM AUTHOR]

It's comparatively difficult to model the solar radiation pressure when calculating the precise orbit of the spacecraft. Especially when the shape of the spacecraft is involute, there would be superposition and obscuration between different materials, so that the normal experimental method doesn't work. Based on the precise 3-D Geometric Model, the simulative vision was used to get the section shape of the spacecraft which perpendicular to the Sun vector, and then calculated the equivalent SRP area and the coefficient with the SRP characteristics of the material. To reduce the amount of calculation, a new method was provided using prior-calculating and real-time-interpolation. Considering that the SRP characteristics would change in space, the calculation was transformed into a semi-empirical formula so the coefficients could be estimated by the observation data. The simulation indicates that the analysis method can effectively reflect the obscuration between the parts of the spacecraft. [ABSTRACT FROM AUTHOR]

It is found that coated spacecraft interfaces separate unsuccessfully despite without cold welding during the spacecraft manufacturing. To investigate the mechanism, effective factors and protection scheme of the phenomenon an experiment was carried on. In this experiment, the aluminium alloys blocks which were coated with nothing, colorless oxidation, MoS2, bright gold plating, and white anodic film respectively were pressed together in pairs, and placed in a high vacuum tank immovably for a period of time. And the separating force between the blocks was tested outside the immovably high vacuum tank. If the experimental blocks were separated unsuccessfully, new copies were placed immovably in the natural pressure and several different temperatures for a period of time. Then the separating force was tested, and the shape changing and element transfering conditions of the interfaces were analyzed. The experiment result shows that there s much tiny concavo-convex grain in the interface of the blocks. Because of the interlocked grain, the blocks have separated unsuccessfully. Hardness and roughness of coating is the immediate cause of this phenomenon, which is called physical tabling. For immovable spacecraft interfaces, a protection method that the coating should be harder and smoother was proposed and validated. [ABSTRACT FROM AUTHOR]

It is very necessary to execute orbit correction for ultra-low-orbit vehicle, which is badly influenced by near-earth environment. The orbital motion equation which considered the non-spherical perturbation and atmospheric perturbation was linearized. Then, an impulse correction method was introduced and used to keep a near-earth vehicle flying along the standard orbit. Furthermore, a method to eliminate errors caused by linearization was also studied. The feasibility of the correction method was validated by simulation. For a 120km-altitude vehicle, flying one hour, the optimal correction times is 6, the required Av is 63. 797 1 m/s. [ABSTRACT FROM AUTHOR]

A new concept of dynamic reentry footprint was presented to ensure the reentry of the spacecraft safely and accurately. Concerned with all the given flight states and the constraint conditions of the spacecraft during the whole reentry flight process, the maximum reachable area on the earth surface can be calculated by the given numerical algorithm. Combined with the dynamic reentry numerical algorithm, the modified genetic algorithm was applied to optimize the reentry flight trajectories, and all the qualified trajectories which constitute the landing footprint can be calculated. Through a series of reentry flight numerical simulations, the characteristics of dynamic reentry footprint was summarized, which can be used to optimize the guidance and flight trajectory. [ABSTRACT FROM AUTHOR]

The guidance law for autonomous rendezvous and proximity operations requires the real-time guidance capability and excellent robustness. An iterative guidance methodology based on convex optimization with a target satellite in an arbitrary orbit was presented. Firstly, by the lossless simplification and convex processing to the dynamic model of relative motion, a convex optimization problem was formed, and its solution was proven to be equivalent to that of the original autonomous rendezvous problem. Secondly, after solving the convex optimization problem by the primal-dual interior point algorithm, an iterative guidance method was used to finish the onboard planning of the chaser vehicle. Numerical experiments demonstrate the efficiency of the proposed guidance method. [ABSTRACT FROM AUTHOR]

According to the characteristics of the magnetic field distribution in the earth space, the propulsion technology without propellant consumption was proposed to maintain or change the LEO spacecraft altitude. The magnetic thrust dynamics was modeled based on the fundamental magnetics, the magnetism propulsion principle with continuous thrust hut no propellant consumption was clarified, and the tracking control strategy of geomagnetic line of force for decoupling the magnetic moment was proposed. The flight envelope of magnetism propulsion and typical calculations for the orbit altitude maintenance or changes was carried out. The results show that the magnetic moment of the spacecraft can be used to maintain or change the spacecraft altitude from 600 km to 1000 km in low earth orbit (LEO) when it is up to 106 Am². Furthermore, the technical feasibility of obtaining high magnetic moment is briefly analyzed. [ABSTRACT FROM AUTHOR]

For the final phase of spacecraft autonomous rendezvous and docking, the vision relative navigation method based on the information of the specific reference light sources was presented. Under the assumption of small attitude angle, the relative navigation models were deduced. The closed-form solution for relative attitude and position parameters based on the vision navigation models was derived. The reduced order Unscented Kalman Filtering (UKF) algorithm was designed to estimate the information of relative attitude angle and relative attitude angle velocity. Simulation results show that the proposed approach provides a convenient way for obtaining the information of relative position and relative attitude, and the reduced order UKF can improve the accuracy of the relative attitude determination. [ABSTRACT FROM AUTHOR]

The traditional optimization algorithms might fail to get the global optimal solution for high non-linear characteristics in engineering optimization of multi-structures. So the Quantum Particle Swarm Optimization (QPSO)algorithm was induced and the QPSO algorithm based on surrogated model was studied. The applications to the structural parameter optimization of the ring truss deployable antennas were carried out by the surrogated model of dynamic characters for ring truss deployable antennas. The results show that the computation efficiency of QPSO algorithm based on Kriging model is acceptable. The non-linear characteristics of the 1st natural frequency and maximum impulse response of the ring truss deployable antennas are remarkable. Then the computation accuracy is validated by comparing the solutions between the optimal result and the finite element result. It's proved that QPSO algorithm based on Kriging model can be used to optimize the complicated structures in engineering design. [ABSTRACT FROM AUTHOR]

The characteristics and effects of analog CMOS integrated circuit on spacecraft were analyzed in the radiation environment. Based on the generation of radiation effect, the main anti-radiation design methods were introduced for the analog CMOS integrated circuit designing and processing. In the outer space, threshold voltage deviation, transdiode decreasing, substrate leakage current increasing and corner noise amplitude increasing occur to the CMOS semiconductor components in the analog CMOS integrated circuit. As a result, there are three kinds of methods proposed to protect against radiation of the analog integrated circuits, including anti-radiation analog CMOS integrated circuit', anti-radiation PCB' and silicon on insulator anti-radiation processing. Accordingly, the designed anti-radiation analog CMOS integrated circuits obtain the ideal effect in anti-radiation function. [ABSTRACT FROM AUTHOR]

As electric power supply devices of spacecraft, solar arrays have played an important role in assuring successful spaceflight. Based on the analysis of some characteristic parameters of solar array deployment process, the equivalent driving moment obtained from every hinge driving moment measurement value was selected as reliability characteristic parameter of solar array deployment. By stress-strength interference theory, a reliability assessment method was put forward which appling hinge measurement data of both driving moment and resistance moment to reliability assessment of the solar array deployment. This method provides a technical approach to validate the reliability of spacecraft solar array deployment. [ABSTRACT FROM AUTHOR]

A precise dynamic model of a novel attitude control actuator--the fluidic ring actuator (FRA) was derived. Based on the practical factors of the flow in flow field, the variation of velocity caused by viscous was analyzed. The equivalent angular momentum of this novel actuator was deduced by integrating the velocity on every cross-section under the consideration of fluid properties. Then, the attitude dynamics of the spacecraft with three orthogonal FRAs was established. An attitude control law was designed according to the attitude dynamics of the spacecraft-FRAs system. Finally, simulation results demonstrate that the output torque has such advantages as fast reaction rate and small error by using the precise model of FRA. The application of the precise model of FRA makes the attitude control of a spacecraft more efficient and accurate. [ABSTRACT FROM AUTHOR]

Traditional formation keeping orbit design is based on linear relative motion model which will lead to more fuel consumption during long-term spacecraft formation flying. Firstly, the periodic condition and periodic solution were obtained by solving the elliptical orbit relative motion model considering the second-order nonlinear terms with perturbation approach. The periodic solution obtained was then used as the reference orbit of a PD control law for formation keeping based on the Lyapunov stability theory. Simulation results show that the periodic solution based on the nonlinear relative motion model of elliptical orbit is more accurate than that based on linear relative motion model of elliptical orbit. And the formation maintain control law with periodic reference orbit based on nonlinear model can save much more fuel compared with that based on the linear model. [ABSTRACT FROM AUTHOR]

In order to improve the performance in singularity avoidance and escape of the dynamic steering law for single gimbal control moment gyroscopes (SGCMGs), a hybrid steering law with feedback and feedforward was proposed. When SGCMGs was not singular, the dynamic steering law with feedback was used. When SGCMGs approaching to or locked in the singular points, the feedforward was constructed to help them avoid or detour round the singular points. Compared with the conventional dynamic steering laws, the hybrid one has much more superior performances in singularity avoidance and escape. Meanwhile, the algorithm has little calculation amount and is easy to implement since it only needs the transpose rather than the pseudo inverse of the Jacobian matrix. Extensive simulation results demonstrate the effectiveness of the algorithm. [ABSTRACT FROM AUTHOR]

The space manipulator joint has to adopt the high reduction rate gearbox. The FEM showed its disadvantage in the stiffness analysis and joint weight reduction due to the complexity of the gearbox. The stiffness model was established with the lumped parameter method and the stiffness series principle. The influence of all parts' stiffness on the general stiffness of the space manipulator joint was analyzed and a new way to optimize the high-reduction-rate joint was given. [ABSTRACT FROM AUTHOR]

Aiming at the large initial error and the strong nonlinear problem existed in Spacecraft attitude determination, an improved regularized auxiliary particle filter (IRAPF) algorithm was proposed. The regularized Particle Filter (RPF) was linked with the auxiliary particle filter (APF), and the fast Gauss transform (FGT) was introduced into the algorithm to improve convergence efficiency. This algorithm can restrain particle degenerating, optimizing sampling by using the newest observed particles. By introducing FGT, the calculation was reduced 30% compared with RAPF method and the real time feature was improved. Simulation results prove the validity of the algorithm. [ABSTRACT FROM AUTHOR]

The traditional angle measurement methods of microwave radar are difficult to satisfy accuracy requirements under the rendezvous and docking near-field conditions. To achieve the high-precision angle measuring in a wide range including near-field as well as far-field, a new algorithm based on phase difference complex vector matching was proposed. Firstly, an object function was defined to quantify the similarity of the phase difference complex vector. Secondly, a nonlinear programming model was established to estimate the azimuth and elevation angles and the solving steps of the model were described in detail. Finally, the results of numerical simulation, airborne flight test and anechoic chamber experiment prove the effectiveness of the algorithm with a high precision of 0.12°. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

In order to describe the relative position and attitude measurement information of RVD spacecraft, a combination method of binocular vision measurement and Plucker linear equation was proposed to determine the relative position and attitude parameters of RVD spacecraft. The relative position and attitude of spacecraft was separately measured and calculated in traditional algorithm, while, the Plucker linear equation jointly describes relative position and attitude information of the two spacecraft. At first, the binocular vision algorithm was used to calculate two non-coplanar lines' coordinate value in tracking spacecraft. Then, the two lines' relative position and attitude relation in two coordinates calculated by Plucker linear equation was obtained. Finally, the relative position and attitude information of the two spacecraft was calculated by the singular value decomposition method. Simulation results show that the algorithm can not only jointly measure position and attitude, but also meet the measurement requirements of RVD spacecraft and verify the scientific rationality and calculation efficiency of the algorithm. [ABSTRACT FROM AUTHOR]

The determination of the overall technology index is a key to the overall design of space manipulator. An overall performance indexes determination method of a space manipulator were presented. The dynamic simulation study was also performed. Firstly, considering the mission objectives and combining with the overall constraints, the main technology indexes, including manipulator length, end-effector positing and orientating accuracies, the maximum end-effector velocities, and joint drive torques, and so on, were demonstrated. Secondly, the multi-body dynamic model of a space robotic system was created. Based on this model, the manipulation process of the space robot carrying a payload under normal and extreme conditions was simulated. Simulation results show that the proposed ideas and methods are very useful for the development of the space robot system. [ABSTRACT FROM AUTHOR]

The configuration efficiency and controllable efficiency are important indices in evaluating the configuration of a single gimbal control moment gyroscope (SGCMG) system. But the configuration efficiency is non-normalized, and the controllable efficiency depends on the magnitude of the angular momentum envelop. By normalizing the configuration efficiency and excluding the impact of the angular momentum envelop to the controllable efficiency, two improved indices, which are capacity ratio and nonsingular capacity ratio, were introduced. As the singularity of SGCMGs is complex and these scalar indices do not indicate the whole properties of the configuration, visualization techniques were used to compare the singular surfaces of different configurations. The analysis results show that the quality of the SGCMG and the symmetries of the configurations mainly affect the configuration indices and the complexity of the singular surface. [ABSTRACT FROM AUTHOR]

The selection of momentum wheel configuration in satellite attitude control system design was investigated. Five different orientations of momentum wheel system were summarized. The design criteria of momentum wheel reconfiguration, the saturation of the momentum wheel under disturbance torque, energy consumption and reliability in different orientations were proposed by comparing the angular momentum envelope. The proposed method could be applied to the design of momentum wheel control system. [ABSTRACT FROM AUTHOR]