نشریه علوم و فناوری فضایی
سال پنجم شماره 4 (پیاپی 13، زمستان 1391)

Star tracker is an attitude determination device which determines the satellite or spacecraft’s attitude using the star position information in inertial and body references. Star information is collected and stored onboard as a “Star or mission catalog”. There are several star catalogs that contain different kinds of information with different accuracy. In this paper the most used star catalogs are introduced and a few star catalog selection features are recommended. These features are weighted according to the star tracker mission type. For the selected star tracker mission, results demonstrate that Hipparcos star catalog is the best choice. Eventually using Hipparcos star catalog, a mission catalog is developed to be used onboard a typical star tracker.

Satellite software is a kind of application that are exposed to soft-errors or transient faults because they work in an environment, which is full of radiations. This kind of fault has caused error in the execution of software. Satellite designers have used different methods, such as hardware shielding and component hardening, for removing the effects of this type of errors. These methods will increase the cost, weight and power consumption. Also, thesemethods decrease the performance. Recently, researches have shown that instead of spatial components,, COTS components can be used in LEO satellites. In addition, software implemented hardware fault tolerance (SIHFT) techniques can be added to these components for dealing with transient faults. In this paper, a branch of these techniques, which is called control flow based techniques, are investigated and the useful method among them is highlighted. The result of the implementation, testing and evaluation of the selected method (i.e. RSCFC) is also presented in this paper.

he application of the hydrogen peroxide as the oxidizer of a HTPB-based prototype ybrid motor is investigated in this paper. First of all, by studying the thermochemical haracteristics of the propellant,operation range of different compositions is defined. A rototype hybrid motor is then designed and manufactured, which is able to produce 10 kg propellant thrust. Conducting a special procedure available stabilized commercial hydrogen peroxide is processed and converted to a 90% density propellant which has less than 15 ppm tainting. A catalytic igniter is used in order to fire the motor. It was observed that the approach of employing the oxidizer to the catalyst bed is highly effective on the quality of the system performance. So, the injection quality of the oxidizer was investigated for different mass flow rates. The performance of the hybrid motor by HTTPB fuel is studied in a successful experimental test. The results of the test including the pressure and characteristic velocity are compared with the predicted theoretical simulations. Moreover, the performance specifications of the motor like the regression rate are determined and compared with the similar researches.

One of the most important issues in the productivity of the satellites is the mode timing and the automatic writing of the scenario on-board of the satellite with regard environmental conditions. In order to have the capability to make flight scenarios for missions sent from ground stations or missions that are generated automatically, on board timing forecastalong with the time to reach the station or the photo target will be necessary. On board automatic generation of the scenarios will reduce the dependency of the satellite to ground station. Calculation of the satellite rise and set times, without the satellite passes and orbit disturbances simulation, would not be possible.This paper presents formula for the satellite passes simulation, algorithm and time prediction computation modeling using keplerian orbital parameters and reduces the size of the calculation utilizing of semimajor axis of the orbit to be used on the board of satellite. To reduce the volume of calculations, it will be necessary to use numerical methods for finding function minimums or roots. semimajor axis of the orbit will decrease the numerical computation to a great extent using the algorithm presented in this paper.

In this paper we are going to design an attitude control system for a Stereo-Imaging Remote Sensing Satellite using of four pyramidal reaction wheels. In this method, in order to provide the power requirements of the satellite by the energy stored in the reaction wheels, a power management law will be designed and added to the attitude control law to charge and discharge the wheels according to the preset profiles. Therefore, attitude control and power management of the satellite will be carried out simultaneously. In this method, while the satellite is in the light and the batteries are charged, the speed of the wheels increase up to the limited speeds and while the power requirements of the subsystems are inadequate, Power will be returned to the subsystems by reducing the speed of the wheels. The design of this system has been conducted to consider the effects of saturation of the wheels to prevent attitude deviations of satellite while power management is done. Simulation results show the good performance of the designed attitude control and power management system of the satellite.

This paper has tried to control for multiple satellite formation flight under a method called Virtual structure. First the algorithm was implemented in the reference coordinate system has no gradient and then we have developed the same algorithm for a circular orbit around the earth within a set of satellites that are in circulation. Then for each of the satellites we have added collision avoidance controller to the system. And then the control system on the satellites such as SPHERES taking into account the constraints operators has control over their implementation.

Recently, many space missions have been using small satellites, because small satellites are easier and faster to develop and thereby, provide increased launch opportunities. Some of these missions include tasks that required agile maneuvers. In this paper, attitude stability testing of an agile three-degree-freedom micro-satellite simulator – which is equipped with a pyramid arrangement of single-gimbal controlmoment gyros (SGCMGs) – is presented. In the attitude stability testing, the local quadratic regulator (LQR) control strategy is used, which has superiority to other approaches due to its independence of using steering law. This simulator allow to test different control laws by using SGCMGs. In this work, after introducing the actuator and satellite simulator and using the control strategy in the simulator, the attitude stability testing is performed and then, the experimental results are presented and discussed. The results show the attitude stability of the simulator which is exposed to the disturbing toques.

Evaluation of the satellite attitude determination system is very complicate because of need to simulate space environment on the ground. In this paper a laboratory Implementation of attitude estimation using sun sensor and magnetometer is reported.The test setup includes a sun simulator installed in a dark room. The magnetometer measures the local magnetic field. Sensors are rotated using a 2 DOF table. The reference models are adapted with the test setup. Attitude determination using classic and extended kalman filter methods is implemented by sensors data fusion. The test results verify the controlled motion of sensors in the limits of setup accuracy order.