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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.

The purpose of this paper is to investigate the problem of the initial attitude detumbling and acquisition for micro‐satellite using geomagnetism with the aid of the pitch momentum bias, and the application of the feedback linearization method, H_∞ and μ‐synthesize control theory in the robust attitude acquisition controller design.

The pitch flywheels establish the momentum bias state in the beginning of the detumbling stage and keep the momentum bias state thereafter. The geomagnetic change rate feedback detumbling controller is used to detumble the micro‐satellite and the gyroscope rigidity is utilized to capture orbital negative normal orientation in the detumbling and attitude acquisition phase. Feedback linearization method is adopted to obtain the linear attitude dynamics. Based on the feedback linearization model, a quasi proportion differential (PD) controller is designed, meanwhile H_∞ and μ‐synthesis control theories are adopted to synthesis the robust attitude acquisition controllers.

The pitch momentum bias‐aided attitude detumbling and acquisition method make the capture of the orbital negative normal orientation faster and more accurate than the classical initial operation process. Quasi PD and H_∞ have greater robustness than the classical PD attitude acquisition controller in normal geomagnetic case; quasi PD and μ‐synthesis have greater robustness than the classical PD attitude acquisition controller in magnetic storm case.

Provides pitch momentum bias‐aided attitude detumbling and acquisition method for the micro‐satellite and the robust attitude acquisition controller design technology.

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.

This paper seeks to provide an insight into the design and development of the thermal test model (TTM) of X‐Sat, a 120 kg class micro‐satellite, being developed at the Centre. This model was specifically constructed for carrying out a thermal balance test (TBT) in a 4 m diameter vertical thermal vacuum chamber.

The construction of the thermal model followed a structural mock‐up model which was modified thermally to suit the purpose. Specific and careful consideration was given to the geometry and, more importantly, thermal characteristics such as thermal mass, surface properties, etc. to mimic the actual satellite configuration as closely as possible. Test plans were devised to qualify the fabricated components to meet the out‐gassing and other thermal requirements for the model. Design and qualification of supporting frame and linkages for TBT are also covered.

It is possible to simulate the thermal characteristics of a micro‐satellite in orbit under a different mission scenario through proper scaling and using alternative material options while developing TTM.

The paper discusses in detail the simplified cost‐effective approach of constructing TTM and also outlines the various issues to be considered for a TBT. It provides valuable information needed for micro‐satellite designers.

With the advent of micro‐satellites technology, passive thermal controls in the form of surface coatings have become important for onboard thermal management. The thermal coatings, however, suffer outgassing and mass loss due to their direct exposure to harsh thermal environment and high vacuum in space. The purpose of this paper is to discuss testing and evaluation on outgassing of AA6061‐T6 specimen surfaces treated with various types of anodized coatings of different thicknesses and the related mass loss before and after thermal exposure.

Samples of chromic acid, polytetrafluroethylene polymer, and black‐ and brown‐colour anodized aluminum coupons were subjected to high vacuum (∼1×10⁻⁶ mbar), before and after thermal baking at 120°C. Spectrum analysis of the outgassed material to know their quantities and proportion was conducted subsequently using a Quadrupole mass analyzer.

The surface coatings under study complied with the spacecraft requirements for the mass loss of less than 1 percent of the total mass of the coating material used for that surface. The mass spectrum analysis of the outgassed material indicated that the majority of the coating mass loss was on account of water vapours and organic solvents like ethylene.

These results provided a good insight into the reliability of the coating materials studied and the bonding between the aluminum substrates and the coatings.

The coatings and the technology needed for their application on aluminum are readily available. The present work on outgassing and mass loss in a simulated space environment will provide useful insight on their usage for micro‐satellites.

The geomagnetic field vector is a function of the satellite’s position. The position and speed of the satellite can be determined by comparing the geomagnetic field vector measured by on board three-axis magnetometer with the standard value of the international geomagnetic field. The geomagnetic model has the disadvantages of uncertainty, low precision and long-term variability. Therefore, accuracy of autonomous navigation using the magnetometer is low. The purpose of this paper is to use the geomagnetic and sunlight information fusion algorithm to improve the orbit accuracy.

In this paper, an autonomous navigation method for low earth orbit satellite is studied by fusing geomagnetic and solar energy information. The algorithm selects the cosine value of the angle between the solar light vector and the geomagnetic vector, and the geomagnetic field intensity as observation. The Adaptive Unscented Kalman Filter (AUKF) filter is used to estimate the speed and position of the satellite, and the simulation research is carried out. This paper also made the same study using the UKF filter for comparison with the AUKF filter.

The algorithm of adding the sun direction vector information improves the positioning accuracy compared with the simple geomagnetic navigation, and the convergence and stability of the filter are better. The navigation error does not accumulate with time and has engineering application value. It also can be seen that AUKF filtering accuracy is better than UKF filtering accuracy.

Geomagnetic navigation is greatly affected by the accuracy of magnetometer. This paper does not consider the spacecraft’s environmental interference with magnetic sensors.

Magnetometers and solar sensors are common sensors for micro-satellites. Near-Earth satellite orbit has abundant geomagnetic field resources. Therefore, the algorithm will have higher engineering significance in the practical application of low orbit micro-satellites orbit determination.

This paper introduces a satellite autonomous navigation algorithm. The AUKF geomagnetic filter algorithm using sunlight information can obviously improve the navigation accuracy and meet the basic requirements of low orbit small satellite orbit determination.

Nowadays, many micro‐satellites rely on three‐axis stabilization using at least three reaction wheels (RWs) for attitude control. Typically, RWs are mounted separately within the satellite and leads to long wires from each RW to the associated electronics. Placement and other constraints may also arise during integration. Furthermore, the large number of mounting holes required for mounting the RWs directly on the structural panels may weaken the integrity of the structure. This paper highlights a need for a special housing to be designed for mounting three to four RWs as a group at one location for simple and easy three‐axis stabilization of typical micro‐satellites.

Details about the development of the design methodology for such housing are discussed. The basic idea is to place the RWs onto the proposed housing, which will be then mounted on to one of the honeycomb panels used for the satellite structure. Requirements analysis, design, validation, and manufacturing process are covered.

The outcome is a dimensionally optimized housing structure for four RWs with a frequency safety factor of three.

In summary, the designed product satisfied all requirements. In addition, the exercise set out a planned procedure for designing similar housings for other satellite components.

The purpose of this paper is to propose an attitude determination and control scheme for a low‐cost Micro‐satellite with defective inertia. Restricted by the payload design, the z‐axis inertia of this satellite is larger than the x and y axes, which is unstable for natural attitude dynamics.

An original operation mode is designed to avoid z axis from long‐time pointing to the sun during damping, which avoids some unexpected damage. In attitude determination design, EKF and UKF algorithms are compared on estimation accuracy, convergence time and computation complexity in attitude estimation design, which is referred to determine the final estimation scheme. A DSP‐based hardware solution is achieved and a semi‐physical testing and simulation system is built.

Simulation results show the 3‐axis stable mode can be built with the proposed scheme, and the unprotected facet of the satellite can be kept away from long‐time pointing to the sun.

The proposed ADCS scheme can be a reference for the future Micro‐satellite programs which share the similar configuration.