• Aerospace Engineering and Technology
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• v.12 no.2
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• pp.14-23
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• 2013

Fault management design of the satellite describes preparations for failures which can occur during operational phase. Fault management design contains detection and isolation function of anomaly, and also it contains function to maintain the satellite in safe condition until the ground station finds out a cause of failure and takes a countermeasure. Unlike normal operation, safing operation is automatically performed by Power Control and Distribution Unit and Integrated Bus Management Unit which loads Flight Software without intervention of ground station. Since fault management operation is automatical, fault management logic and functionality of relevant hardware should be thoroughly checked during ground test phase, and error which is similar to actual should be carefully applied without damage. Verification test for fault management design is conducted for various subsystems of satellite. In this paper, we show the design process of fault management design verification test for Electrical Power Subsystem and Attitude and Orbit Control Subsystem of Low Earth Orbit satellite flight model and the test results.

• 제어로봇시스템학회:학술대회논문집
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• 2000.10a
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• pp.461-461
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• 2000

Nonlinear pulse width modulation(PWM) controlled system is considered to achieve control performance of thruster-controlled spacecraft. The actual PWM controlled motions occurs, very closely, around the average model rajectory. Furthermore nonlinear PWM controller design can be directly applied to thruster controlled spacecraft to determine thruster on-time. Sliding mode control for attitude tracking of three-axis thruster-controlled spacecraft is presented. Simulation results are shown which use modified Rodrigues parameters and sliding mode control law to achieve attitude tracking of a three-axis spacecraft with thrusters.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.5
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• pp.371-378
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• 2019

High precision pointing satellite uses the reaction wheels for the attitude control and their momentum dumping is performed by the three magnetic torquers. In this paper, the effects of one magnetic torquer's failure on the momentum dumping will be reviewed. When the satellite on the high inclination angle orbit holds LVLH (Local Vertical Local Horizontal) attitude, pitch axis magnetic torquer failure causes the momentum dumping failure. But in case of other torquer's failure, momentum dumping is still possible with degraded dumping performance. When pitch axis magnetic torquer fails, momentum dumping is possible by changing the satellite attitude. This paper propose the satellite attitude change to make the momentum dumping possible when pitch axis magnetic torquer fails. In addition, if torquer arrangement is modified, momentum dumping is always possible regardless of any torquer's failure.

• Journal of Astronomy and Space Sciences
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• v.23 no.2
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• pp.161-166
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• 2006

In this paper, we describe the RLG current stabilizer circuit for attitude control in the satellite. The RLG makes use of the Sagnac effect within a resonant of a HeNe laser. The difference between two discharge currents causes one of the gyro bias errors. The theoretical background and current stabilizer are introduced. It is verified that the circuit designed is applicable to the test of input voltage and temperature.

• 제어로봇시스템학회:학술대회논문집
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• 2000.10a
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• pp.30-30
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• 2000

In the three axis control of satellite by using reaction wheel and gyro, a Gyro carries out measuring of the attitude angie and the attitude angular velocity. The Gyro is operated by the electronic part and the mechanic actuator. The digital part of the electronic part is consisted of the FPGA (Field Programmable Gate Array), which is one of the methods for designing VLSI (Very Large Scale Integrated Circuit), and the mechanic actuator processes the input/output data by the dynamic model. In the research of the mixed mode of Gyro, the simulation is accomplished by SABER of the mixed mode simulator and the results for the practical implementation of the satellite ACS (Attitude Control System) interfaced with the data processing are proposed.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.1
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• pp.50-60
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• 2015

A propulsion system of a satellite provides a necessary thrust to reach to the final orbit after a separation from a launch vehicle. Also, it supplies pulse moments to maintain the satellite in a mission orbit and for its attitude controls during a mission life time. The present study investigates the development trend of bipropellant rocket engines for an orbit transfer and an attitude control of a satellite using monomethylhydrazine and hydrazine for fuel and dinitrogen tetroxide for oxidizer to derive fundamental specifications which are necessary for domestic development researches. Also, their major performance characteristics are summarized.

• Journal of Positioning, Navigation, and Timing
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• v.6 no.4
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• pp.195-203
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• 2017

A 2U cube satellite called SNUGLITE has been developed by GNSS Research Laboratory in Seoul National University. Its main mission is to perform actual operation by mounting dual-frequency global positioning system (GPS) receivers. Its scientific mission aims to observe space environments and collect data. It is essential for a cube satellite to control an Earth-oriented attitude for reliable and successful data transmission and reception. To this end, an attitude estimation and control algorithm, Attitude Determination and Control System (ADCS), has been implemented in the on-board computer (OBC) processor in real time. In this paper, the Extended Kalman Filter (EKF) was employed as the attitude estimation algorithm. For the attitude control technique, the Linear Quadratic Gaussian (LQG) was utilized. The algorithm was verified through the processor in the loop simulation (PILS) procedure. To validate the ADCS algorithm in the ground, the experimental verification via a single axis Hardware-in-the-loop simulation (HILS) was used due to the simplicity and cost effectiveness, rather than using the 3-axis HILS verification (Schwartz et al. 2003) with complex air-bearing mechanism design and high cost.

• ETRI Journal
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• v.25 no.5
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• pp.387-400
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• 2003

Since its launching on 21 December 1999, the Korea Multi-Purpose Satellite-I (KOMPSAT-I) has been successfully operated by the Mission Control Element (MCE), which was developed by the ETRI. Most of the major functions of the MCE have been successfully demonstrated and verified during the three years of the mission life of the satellite. This paper presents the operational performances of the various functions in MAPS. We show the performance and analysis of orbit determinations using ground-based tracking data and GPS navigation solutions. We present four instances of the orbit maneuvers that guided the spacecraft form injection orbit into the nominal on-orbit. We include the ground-based attitude determination using telemetry data and the attitude maneuvers for imaging mission. The event prediction, mission scheduling, and command planning functions in MAPS subsequently generate the spacecraft mission operations and command plan. The fuel accounting and the realtime ground track display also support the spacecraft mission operations.

• Journal of Astronomy and Space Sciences
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• v.36 no.3
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• pp.181-186
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• 2019

The attitude information of spacecraft can be obtained by the sensors attached to it using a star tracker, three-axis magnetometer, three-axis gyroscope, and a global positioning signal receiver. By using these sensors, the spacecraft can be maneuvered by actuators that generate torques. In particular, electromagnetic-torque bars can be used for attitude control and as a momentum-canceling instrument. The spacecraft momentum can be created by the current through the electrical circuits and coils. Thus, the current around the electromagnetic-torque bars is a critical factor for precisely controlling the spacecraft. In connection with these concerns, a solar-cell array can be considered to prevent generation of a magnetic dipole moment because the solar-cell array can introduce a large amount of current through the electrical wires. The maximum value of a magnetic dipole moment that cannot affect precise control is $0.25A{\cdot}m^2$, which takes into account the current that flows through the reaction-wheel assembly and the magnetic-torque current. In this study, we designed a 300-W solar cell array and presented an optimal wire-routing method to minimize the magnetic dipole moment for space applications. We verified our proposed method by simulation.

• Journal of Power Electronics
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• v.9 no.4
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• pp.631-642
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• 2009

The sun is a useful reference direction because of its brightness relative to other astronomical objects and its relatively small apparent radius as viewed by spacecrafts near the Earth. Most satellites use solar power as a source of energy, and so need to make sure that solar panels are oriented correctly with respect to the sun. Also, some satellites have sensitive instruments that must not be exposed to direct sunlight. For all these reasons, sun sensors are important components in spacecraft attitude determination and control systems. To minimize components and structural mass, some components have multiple purposes. The solar cells will provide power and also be used as coarse sun sensors. A coarse Sun sensor is a low-cost attitude determination sensor suitable for a wide range of space missions. The sensor measures the sun angle in two orthogonal axes. The Sun sensor measures the sun angle in both azimuth and elevation. This paper presents the development of a model to determine the attitude of a small cube-shaped satellite in space relative to the sun's direction. This sensor helps small cube-shaped Pico satellites to perform accurate attitude determination without requiring additional hardware.