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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.

This paper aims to present an attitude determination and control system for a nano‐astrometry satellite which requires precise angular rate control. Focus of the research is methods to achieve the requirement.

In order to obtain astrometry data, the satellite attitude should be controlled to an accuracy of 0.05°. Furthermore, attitude spin rate must be controlled to an accuracy of 4×10⁻⁷ rad/s during observation. In this paper the following unique ideas to achieve these requirements are introduced: magnetic disturbance compensation and rate estimation using star blurred images.

This paper presents the feasibility of a high accurate attitude control system in nano‐ and micro‐satellite missions.

This paper presents a possibility of the application of nano‐satellites to remote‐sensing and astronomy mission, which requires accurate attitude control.

Originalities of the paper are the methods to achieve the high accurate attitude control: magnetic disturbance compensation and angular rate estimation using star images.

Fractionated satellite clusters need coming together and avoiding crashing with limited random initial relative motion conditions. It is not necessary to keep the invariable configuration. The purpose of this paper is to put forward a control law which simulates organism swarm motion.

The motion of satellites follows three rules: coming together, velocity homology and avoiding crash. According to the rules, three control forces should be applied to satellite individuals. The final control force is the sum of three control forces. Electromagnetic dipole strengths calculation is formulated as nonlinear optimization problem. Considering control strengths have to be got in real time, iterative steps of optimization algorithm are fixed.

A control law which simulates organism swarm motion is put forward. The simulation shows the organism swarm motion simulation control law can keep fractionated satellite cluster coming together and avoiding crash. When iterative steps of optimization algorithm is fixed, the error of solve nonlinear equations is acceptable.

The control law is robust and easy to realize. When electromagnetic satellite cluster need not keep fixed configuration, it is a choice of control law of relative motion.

The purpose of this paper is to provide a feasible method to solve the zenith pass problem that can occur when the inter‐satellite linkage antenna of the user satellite is tracing TDRS. The antenna uses the elevation‐over‐azimuth architecture.

The movement laws of the inter‐satellite linkage can be obtained based on the orbit predictions of the user satellite and TDRS. According to the movement laws, the zenith pass moments and blindness zones are found. The trajectory preprocessor is provided to design a command trajectory for driving the axis of the tilting mechanism.

In the worst situation, the blindness zone can appear once every half day. Three special orbit altitude values are obtained. When the user satellite picks one of them as its orbit altitude, the blindness zone may be avoided forever. The zenith pass tracing strategies based on the mechanical tilting method have been designed.

This method obtains the stable tracking during the zenith pass course by changing the hardware structure of the antenna. It is too expensive and can influence the pointing precision of the antenna.

The research can help the engineers analyze and solve the zenith pass problem of the antenna.

This paper studies the zenith pass problem that can occur when the inter‐satellite linkage antenna of the user satellite is tracing TDRS and provides a solving method.

The purpose of this paper is to propose a fuel‐optimal virtual centre selection method for formation flying maintenance in the J2 perturbed environment.

Based on the relative orbital elements (ROE) theory, the J2 perturbed relative motions between different satellites in the formation are analyzed, and then the fuel‐optimal virtual centre selection issue for formation flying maintenance are parameterized in terms of ROE. In order to determine the optimal virtual centre, two theories are proposed in terms of ROE.

Numerical simulations demonstrate that the fuel‐optimal virtual centre selection method is valid, and the control of the ROE of each satellite with respect to a virtual optimal centre of the formation is more efficient regarding the fuel consumption than the control of all satellites with respect to a satellite belonging to the formation.

The fuel‐optimal virtual centre selection method is valid for formation flying mission whose member satellite in circular or near circular orbit.

The fuel‐optimal virtual centre selection approach can be used to solve formation flying maintenance problem which involves multiple satellites in the formation.

The paper proposes a fuel‐optimal virtual centre selection method in terms of ROE, and shows that keeping the formation with respect the optimal virtual centre is more fuel efficient.

The purpose of this study is to address a highly heterogeneous rift margin environment and exhibit considerable spatiotemporal hydro-climatic variations. In spite of limited, random and inaccurate data retrieved from rainfall gauging stations, the recent advancement of satellite rainfall estimate (SRE) has provided promising alternatives over such remote areas. The aim of this research is to take advantage of the technologies through performance evaluation of the SREs against ground-based-gauge rainfall data sets by incorporating its applicability in calibrating hydrological models.

Selected multi satellite-based rainfall estimates were primarily compared statistically with rain gauge observations using a point-to-pixel approach at different time scales (daily and seasonal). The continuous and categorical indices are used to evaluate the performance of SRE. The simple scaling time-variant bias correction method was further applied to remove the systematic error in satellite rainfall estimates before being used as input for a semi-distributed hydrologic engineering center's hydraulic modeling system (HEC-HMS). Runoff calibration and validation were conducted for consecutive periods ranging from 1999–2010 to 2011–2015, respectively.

The spatial patterns retrieved from climate hazards group infrared precipitation with stations (CHIRPS), multi-source weighted-ensemble precipitation (MSWEP) and tropical rainfall measuring mission (TRMM) rainfall estimates are more or less comparably underestimate the ground-based gauge observation at daily and seasonal scales. In comparison to the others, MSWEP has the best probability of detection followed by TRMM at all observation stations whereas CHIRPS performs the least in the study area. Accordingly, the relative calibration performance of the hydrological model (HEC-HMS) using ground-based gauge observation (Nash and Sutcliffe efficiency criteria [NSE] = 0.71; R² = 0.72) is better as compared to MSWEP (NSE = 0.69; R² = 0.7), TRMM (NSE = 0.67, R² = 0.68) and CHIRPS (NSE = 0.58 and R² = 0.62).

Calibration of hydrological model using the satellite rainfall estimate products have promising results. The results also suggest that products can be a potential alternative source of data sparse complex rift margin having heterogeneous characteristics for various water resource related applications in the study area.

This research is an original work that focuses on all three satellite rainfall estimates forced simulations displaying substantially improved performance after bias correction and recalibration.

X-ray pulsar navigation (XPNAV) is an autonomous celestial navigation technology for deep space missions. The error in the pulse time of arrival used in pulsar navigation is large for various practical reasons and thus greatly reduces the navigation accuracy of spacecraft near the Earth and in deep space. This paper aims to propose a novel method based on ranging information that improves the performance of XPNAV.

This method replaces one pulsar observation with a satellite observation. The ranging information is the difference between the absolute distance of the satellite relative to the spacecraft and the estimated distance of the satellite relative to the spacecraft. The proposed method improves the accuracy of XPNAV by combining the ranging information with the observation data of two pulsars.

The simulation results show that the proposed method greatly improves the XPNAV accuracy by 70% compared with the conventional navigation method that combines the observations of three pulsars. This research also shows that a larger angle between the orbital plane of the satellite and that of the spacecraft provides higher navigation accuracy. In addition, a greater orbital altitude difference implies higher navigation accuracy. The position error and ranging error of the satellite have approximately linear relationships with the navigation accuracy.

The novelty of this study is that the satellite ranging information is integrated into the pulsar navigation by using mathematical geometry.

This paper aims to describe a novel type of attitude control system (ACS) in different configurations. This servomechanism is compared with control moment gyro (CMG) in significant parameters of performance for ACS of rigid satellite.

This new actuator is the fluid containing one or more rings and fluid flow is supplied by pump. The required torque control is obtained by managing fluid angular velocity. The cube-shaped satellite with three rings of fluid in the principle axes is considered for modeling. The satellite is considered rigid and nonlinear dynamics equation is used for it. In addition, the failure of the pyramid-shaped satellite with an additional ring fluid is discussed.

The controller model for four fluid rings has more complexity than for three fluid rings. The simulation results illustrated that four fluid rings need less energy for stabilization than three fluid rings. The performance of this type of actuator is compared with CMG. At last, it is demonstrated that performance parameters are improved with fluid ring actuator.

Fluid ring actuator can be affected by environmental pressure and temperature. Therefore, freezing and boiling temperature of the fluid should be considered in system designation.

Fluid ring servomechanism can be used as ACS in rigid satellites. This actuator is compared by CMG, the prevalent actuator. It has less displacement attitude maneuver.

The results provide the feasibility and advantages of using fluid rings as satellite ACS. The quaternion error controller is used for this model to enhance its performance.

The purpose of this paper is to present a method to conduct small satellite rendezvous mission by using the differential aerodynamic forces under J2 perturbation in low earth orbit (LEO).

Each spacecraft is assumed to be equipped with two large flat plates, which can be controlled for generating differential accelerations in all three directions. Based on the kinetic theory, the aerodynamic lift and drag generated by a flat plate are calculated. To describe the relative dynamics under J2 perturbation, a modified model is derived from the high-fidelity linearized J2 equations proposed by Schweighart and Sedwick.

Simulation results demonstrate that the proposed method is valid and efficient to solve satellite rendezvous problem, and the modified model considering J2 effect shows better accuracy than the Horsley’s Clohessy–Wiltshire-based model.

Because aerodynamic force will reduce drastically as orbital altitude rises, the rendezvous control strategy for small satellites presented in this paper should be limited to the scenarios when satellites are in LEO.

The rendezvous control method in this paper can be applied to solve satellite rendezvous maneuver problem for small satellites in LEO.

This paper proposes a modified differential aerodynamic control model by considering J2 perturbation, and simulation results show that it can achieve higher rendezvous control accuracy.

This paper aims to provide a new method to design a fuel efficient control strategy such as J₂ perturbation for deploying a constellation into a specified configuration. The nonspherical perturbation, mainly J₂ perturbation, is the dominant perturbation for low-Earth-orbit (LEO) satellites of a constellation. This perturbation can be utilized in the control strategy to lower fuel consumption enormously.

The relationship of the coupled variables, the relative right ascension of ascending node (RAAN) and the relative phase (RP), are analyzed. First-order approximation expressions of the relative RAAN (RRAAN) and the relative phase (RP) with respect to the semimajor axis are derived. According to the Gauss’ variational equations, the reduced explicit functions of these variables in regard to each active control are established. Based on these functions, control strategy design methods, including the preliminary planning and iterative corrections, are proposed. The numerical simulation is carried out to verify the proposed method.

The results indicate that the constellation can be deployed accurately about the semimajor axis, the RRAAN and the relative phase (RP) by the developed fuel efficient control strategy.

The proposed control strategy is limited for the orbital altitude where the J₂ perturbation is dominant.

The proposed effective method is applicable for the engineers planning an orbital control strategy of deploying satellites of a constellation.

The new control strategy can realize utilization of J₂ perturbation and an accurate deployment, simultaneously. Further, this paper provides practical help for satellite engineers.