• Bulletin of the Korean Space Science Society
• /
• 2010.04a
• /
• pp.27.5-28
• /
• 2010

The unexpected observation condition or insufficient measurement modeling can lead to uncertain measurement errors. The uncertain measurement error of orbit determination problem typically consists of noise, bias and drift. It must be removed by using a proper estimation process for better orbit accuracy. The estimation of noise and drift is not easy because of their random or unpredictable variation. On the other hand, bias is a constant difference between the mean of the measured values and the true value, so it can be simply removed. In this study, precision orbit determination with SLR observations considering range bias estimation is presented. The Yonsei Laser-ranging Precision Orbit Determination System (YLPODS) and SLR NP (Normal Point) observations of CHAMP satellite are used for this work. The SLR residual test is performed to estimate the range bias of each arc. The result shows that we can get better orbit accuracy through range bias estimation.

• 제어로봇시스템학회:학술대회논문집
• /
• 1997.10a
• /
• pp.155-158
• /
• 1997

In this paper, we present new results concerning the relationship between the input-output and Lyapunov stability of nonlinear system possessing a periodic orbit. Definition of small-signal finite-gain L$\sub$p/ stability around periodic orbit is introduced. We show L$\sub$p/ stability of exponentially stable periodic orbit using quadratic Lyapunov functions for the periodic orbit. The L$\sub$2/ gain analysis is presented with Hamiltonian-Jacobi inequality along with an example.

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

• Bulletin of the Korean Space Science Society
• /
• 2011.04a
• /
• pp.24.4-24.4
• /
• 2011

Frozen orbit concept is very useful in designing particular mission orbits including the Sun-synchronous and minimum altitude variation orbits. In this work, variety of frozen and Sun-synchronous orbit conditions around the Moon is investigated and analyzed. The first two zonal harmonics of the Moon, J2 and J3, are considered to determine mean orbital elements to be a frozen orbit. To check the long-term behavior of a frozen orbit, formerly developed YonSei Precise Lunar Orbit Propagator (YSPLOP) is used. First, frozen orbit solutions without conditions to be the Sun-synchronous orbit is investigated. Various mean semi-major axes having between ranges from 1,788 km to 1,938 km with inclinations from 30 deg to 150 deg are considered. It is found that a polar orbit (90 deg of inclination) having 100 km of altitude requires the orbital eccentricity of about 0.01975 for a frozen orbit. Also, mean apolune and perilune altitudes for this case is about 136.301 km and 63.694 km, respectively. Second, frozen orbit solutions with additional condition to be the Sun-synchronous orbit is investigated. It is discovered that orbital inclinations are increased from 138.223 deg to 171.553 deg when mean altitude ranged from 50 km to 200 km. For the most usual mission altitude at the Moon (100 km), the Sun-synchronous orbit condition is satisfied with the eccentricity of 0.01124 and 145.235 deg of inclination. For this case, mean apolune and perilune altitudes are found to be about 120.677 km and 79.323 km, respectively. The results analyzed in this work could be useful to design a preliminary mapping orbit as well as to estimate basic on-board payloads' system requirements, for a future Korea's lunar orbiter mission. Other detailed perturbative effects should be considered in the further study, to analyze more accurate frozen orbit conditions at the Moon.

• Journal of Astronomy and Space Sciences
• /
• v.15 no.1
• /
• pp.235-244
• /
• 1998

The solutions of orbital maneuver problem using the sliding mode control in the presence of the erath gravitational perturbations is obtained. Especially, the optimization of consuming fuel for maneuver is performed. The impulsive solution to Lambert's problem using the combined equation method to minimize total ${\Delta}V is used for the desired orbit and the maneuver times. Two-step sliding mode control method is introduced for satisfying the boundary conditions of finite-thrust rendezvous problem at the end of maneuver time. Using the new approach to the orbit maneuver problem, two-step sliding mode control, orbit maneuvers are processed. The solutions to a rendezvous using the optimal control are obtained, and they are compared to the results by two-step sliding control.According to the new approach for orbit maneuver, the thrust-coast-thrust type controller is obtained to make satellite to track desired Lambert's orbit, and the total ${\Delta}V$ required for maneuver is resonable in comparison with the impulsive solution to Lambert's problem. The final state variables, also are close to the boundary conditions at the end of maneuver times.

• 제어로봇시스템학회:학술대회논문집
• /
• 1994.10a
• /
• pp.218-223
• /
• 1994

On the satellite communication system, conventional key issues of control have been focused on the attitude and orbit control, monitoring and control of communication payload such as IOT(In-Orbit-Test) and CSM(Communication System Monitoring) and so on. As the vulnerabilities are being increased on the satellite communication network, security services are required to protect it against security violated attacks. In this paper, a security architecture for satellite communication network is presented in order to provide security services and mechanisms. Authentication protocol and encryption scheme are also proposed for spacecraft command authentication and confidentiality.

• Journal of Astronomy and Space Sciences
• /
• v.33 no.1
• /
• pp.37-44
• /
• 2016

This paper presents an algorithm for Real-Time Orbit Determination (RTOD) of navigation satellites for the Korean Regional Navigation Satellite System (KRNSS), when the navigation satellites generate ephemeris by themselves in abnormal situations. The KRNSS is an independent Regional Navigation Satellite System (RNSS) that is currently within the basic/preliminary research phase, which is intended to provide a satellite navigation service for South Korea and neighboring countries. Its candidate constellation comprises three geostationary and four elliptical inclined geosynchronous orbit satellites. Relative distance ranging between the KRNSS satellites based on Inter-Satellite Ranging (ISR) is adopted as the observation model. The extended Kalman filter is used for real-time estimation, which includes fine-tuning the covariance, measurement noise, and process noise matrices. Simulation results show that ISR precision of 0.3-0.7 m, ranging capability of 65,000 km, and observation intervals of less than 20 min are required to accomplish RTOD accuracy to within 1 m. Furthermore, close correlation is confirmed between the dilution of precision and RTOD accuracy.

• Journal of Aerospace System Engineering
• /
• v.7 no.3
• /
• pp.26-32
• /
• 2013

Passive attitude stabilization method has been widely usde for attitude determination and control of cube satellite due to its advantage of system simplicity. The permanent magnet installed on the cube satellite passively controls the attitude of the satellite such that the satellite is aligned with the earth magnetic field. In this paper, on-orbit thermal behavior of the cube satellite with the permanent magnet attitude stabilization method has been investigated through on-orbit thermal analysis. THe orbit profile obtained from the aforementioned attitude control method has been reflected in the analysis. The analysis results indicate that the thermal design proposed in this study is effective for satisfying the temperature requirements of the commericial mission equipments.

• Journal of Astronomy and Space Sciences
• /
• v.29 no.3
• /
• pp.275-285
• /
• 2012

In this study, we present preliminary results of precise orbit determination (POD) using satellite laser ranging (SLR) observations for International Laser Ranging Service (ILRS) Associate Analysis Center (AAC). Using SLR normal point observations of LAGEOS-1, LAGEOS-2, ETALON-1, and ETALON-2, the NASA/GSFC GEODYN II software are utilized for POD. Weekly-based orbit determination strategy is applied to process SLR observations and the post-fit residuals check, and external orbit comparison are performed for orbit accuracy assessment. The root mean square (RMS) value of differences between observations and computations after final iteration of estimation process is used for post-fit residuals check. The result of ILRS consolidated prediction format (CPF) is used for external orbit comparison. Additionally, we performed the precision analysis of each ILRS station by post-fit residuals. The post-fit residuals results show that the precisions of the orbits of LAGEOS-1 and LAGEOS-2 are 0.9 and 1.3 cm, and those of ETALON-1 and ETALON-2 are 2.5 and 1.9 cm, respectively. The orbit assessment results by ILRS CPF show that the radial accuracies of LAGEOS-1 and LAGEOS-2 are 4.0 cm and 5.3 cm, and the radial accuracies of ETALON-1 and ETALON-2 are 30.7 cm and 7.2 cm. These results of station precision analysis confirm that the result of this study is reasonable to have implications as preliminary results for administrating ILRS AAC.

• 제어로봇시스템학회:학술대회논문집
• /
• 1991.10a
• /
• pp.425-430
• /
• 1991

For satellite orbit determination, a satellite (K-3H) which is affected by the earth's gravitational field and the earth's atmospheric drag, the sun, and the moon is chosen as a dynamic model. The state vector include orbit parameters, uncertain parameters associated with perturbations and tracking stations. These perturbations include gravitational constant, atmospheric drag, and jonal harmonics due to the earth nonsphericity. Early orbit was obtained with given the predicted orbital parameter of the satellite. And orbit determination, which is applied to Extended Kalman Filter(EKF) for real time implementation , use the observation data which is given by satellite tracking radar system and then orbit estimation is accomplished. As a result, extended sequential estimation algorithm has a fast convergence and also indicate effectiveness for real time operation.