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This paper aims to present a three-axis attitude tracking control law to solve the attitude maneuver of a flexible satellite in the presence of parameter uncertainties and external disturbance.

Based on the relative dynamic equation where the relative attitude is described by quaternion, a robust control law composed of a proportional derivative (PD) part plus a signum function is designed and only requires the measurement of attitude and angular velocity. Furthermore, the stability analysis of the proposed control law is given through a two-step proof technique.

Numerical simulation results demonstrate that fine convergence of the attitude and angular velocity error and low-level vibration of flexible appendages are obtained by the proposed controllers.

The controller with the structure of a PD term plus a switching function about a sliding variable has low computational complexity and does not need to measure the modal variables of elastic appendages, so it can be used in orbit without modification.

The globally asymptotic stability of the controller in the presence of model uncertainties and external disturbances is proven rigorously through a two-step proof technique.

The measurement of geomagnetic field can provide a reliable and economical basis for attitude and orbit information of low earth orbiting satellite. Because the earth's magnetic field is a function of position, and its measurement on the orbit are fully observable, orbit estimation can be obtained using extend Kalman filter (EKF) algorithm. With the assistant of angle velocity information from gyro measurement, attitude estimation can also be obtained. At the same time, gyro drift rate estimation is a part of the filter output. Although orbit and attitude determination are independent of each other, the filter can give the orbit and attitude estimation at the same time. The results of the numerical test show that a signal EKF can estimate both orbit and attitude by using magnetometer and gyro measurement only. The accuracy, usually is sufficient for low earth orbiting satellites.

This article presents the optimized design, analysis and calculation concerning the trajectory of a lunar polar probe. Firstly, the trajectory design principles and constraints are determined. The preliminary design and analysis of the circumlunar orbit, transfer orbit to the moon and earth parking orbit are carried out separately and some computations for the flight trajectory concept have been made too. To reduce the fuel needed for error in orbital maneuver efficiently and satisfy the requirements on the launch window, some detailed design and analysis for the rather advanced phasing earth‐moon transfer orbit are given here and also the strategy for optimum orbit correction.

To provide a new method to determine parameters of the attitude determination system facing micro‐core.

Take example for attitude determination systems based on star‐sensor and fiber‐optic gyroscope combination and only based on star‐sensor. The optimum parameters of sensors are obtained by setting up of optimization design model of the attitude determination system adopting genetic algorithm.

Put forward a new concept of micro‐core aiming at a micro satellite. Further aiming at micro‐core, a new method which differs from traditional satellite design methods is adopted in this paper. The method proposed in this paper is instructive to the design of future micro satellites.

The method proposed in this paper only applied to attitude determination system. With the development of this method, it is hoped that the method can apply to other systems of a micro satellite.

The method proposed in this paper is instructive to the engineering design of a micro satellite.

Put forward a new concept of micro‐core, and aiming at its design a new method is proposed to design the attitude determination system by adopting genetic algorithm. The method is different from traditional satellite design methods.

The HITSAT‐1 is the first small satellite developed by Harbin Institute of Technology (HIT) whose mission objective is to test several pivotal techniques. In the initial orbit period, the satellite is likely to tumble as the result of separating from the rocket. How to capture it promptly with finite magnetic torque is an important problem. In this paper, considering the restrictive conditions of the magnetic field, the variable structure control theory is used to cope with the magnetic torque. Because of using the reaction wheels and magnetorquers as the control actuators, the combination control algorithm has been adopted in the initial orbit period. The results of the computer simulation indicated that the algorithm has excellent robustness and can be designed and realized easily.

The micro‐satellite clusters have been discussed for several years, however, there is not a common framework about its software, and various researches distributed at different domains. In order to conduct the future work well, the purpose of this paper is to systematically describe micro‐satellite clusters' characteristics, clusters software model, and present a distributed testbed to shorten test process, and minimize the development cost.

The cluster characteristics and model is summarized through analyzing the past satellite cluster programs. Then the ground test system is designed to shorten micro‐satellite's development period, improve its reliability.

The clusters' characteristics are discussed, such as coverage, scalability, fault tolerance, low cost, etc. The clusters' data flow and on‐board software architecture are presented according to properties of clusters. Finally, the distributed testbed that focuses on future on‐board software and hardware technologies that aim to rapid design, build, integration, test, deployment, and operation of the future micro‐satellite is designed.

The presentation of software architecture of cluster member can improve the micro‐satellite's development, and the distributed testbed can improve the ground test efficiency, especially, when the micro‐satellite quantity is big.

Combining the characteristic of satellite “minisize nucleus” non‐equilibrium molecular dynamics (NEMD) method is used. We select corresponding Tersoff potential energy function to build model and, respectively, simulate thermal conductivities of silicon nanometer thin film.

NEMD method is used, and the corresponding Tersoff potential energy function is used to build model.

The thermal conductivities of silicon nanometer thin film are markedly below the corresponding thermal conductivities of their crystals under identical temperature. The thermal conductivities are rising with the increase of thickness of thin film; what's more, the conductivities have a linear approximation with thickness of the thin film.

It is difficult to do physics experiment.

The findings have some theory guidance to analyze satellite thermal control.

The calculation results of thermal conductivities specify distinct size effect. The normal direction thick film thermal conductivity of silicon crystal declines with the increasing temperature. The thermal conductivities are rising with the increase of thickness of thin film; what's more, the conductivities have a linear approximation with thickness of the thin film.

To give a new method to calculate the thermal conductivity of thin films which thickness is less than micro‐nanometer when non‐Fourier effect will appear in heat conduction and Fourier law is not applicable for calculating the thermal conductivity.

The Cattaneo equation based on the heat flow relaxation time approximation is used to calculate the thermal conductivity.

The results show that the thermal conductivity is not the thermophysical properties of material, but is the non‐linear function of temperature and film thickness when the dimension of film is less than micro‐nanometer.

The application of this method is limited by little experimental data of heat flow relaxation time for materials other than Ar crystals.

The paper demonstrates how the thermal conductivity of Ar crystals film can be calculated by NEMD algorithm and considers the non‐Fourier effect in the simulation.

This paper sets out to optimize the shape and size of microchannels cooling heat sink, which has been widely used to cool electronic chip for its high heat transfer coefficient and compact structure.

Sequential Quadratic Programming (SQP) method is used to optimize the cross‐section sizes of microchannels. Finite volume method is used to numerically simulate the cooling performance of optimal microchannel cooling heat sink.

The optimized cross‐section shape of microchannel is rectangular, and the width and depth of microchannel is 50 and 1,000 μm, respectively, the number of microchannels is 60, and the corresponding least thermal resistance is 0.115996°C/W. The results show that the heat transfer performance of microchannel cooling heat sink is affected intensively by its cross‐section shape and dimension. The convection heat resistance R_conv between inner surface in microchannels and working fluid has more influence in the total heat resistance. The heat flux of chip is 278 W/cm² and, through the optimization microchannel cooling heat sink, the highest temperature in the chip can be kept below 42°C, which is about half of that without optimizing heat sink and can ensure the stability and reliability of chip.

The convection heat transfer coefficient is calculated approximatively here for convenience, and that may induce some errors.

The optimized microchannels cooling heat sink may satisfy the request for removal of high heat flux in new‐generation chips.

One of the primary problems in the field of on‐orbit service and space conflict is related to the approach to the target. The development of guidance algorithms is one of the main research areas in this field. The objective of this paper is to address the guidance problem for autonomous proximity manoeuvres of a chase‐spacecraft approaching a target spacecraft.

The process of autonomous proximity is divided into three phases: proximity manoeuvre, fly‐around manoeuvre, and final approach. The characteristics of the three phases are analyzed. Considering the time factor of autonomous proximity, different orbits for the three phases are planned. Different guidance algorithms, which are based on multi‐pulse manoeuvres, are then devised.

This paper proposes three phases of autonomous proximity and then designs a guidance method, which hinges on a multi‐pulse algorithm and different orbits for the three phases; in addition, a method of impulse selection is devised.

An easy methodology for the analysis and design of autonomous proximity manoeuvres is proposed, which could also be considered for other space applications such as formation flying deployment and reconfiguration.

Based on this guidance method, the manoeuvre‐flight period of the chase‐spacecraft can be set in accordance with the mission requirements; the constraints on fuel mass and manoeuvre time are both considered and satisfied. Consequently, this proposed guidance method can effectively deal with the problem of proximity approach to a target spacecraft.