Bang-Yeop Kim

Geostationary orbit determination for time synchronization using analytical dynamic models

A real time analytical orbit determination method has been developed for precision national time synchronization. The one-way time transfer technique via a geostationary TV satellite standard time and frequency signal (STFS) dissemination system was considered. The differential method was also applied for mitigating errors in geostationary satellite STFS dissemination system. Analytical dynamic orbit determination with extended Kalman filter (EKF) was implemented to improve differential mode STFS (DSTFS) service accuracy by acquiring better accuracy of a geostationary satellite position. The perturbation force models applied for satellite dynamics include the geopotential perturbation up to fifth degree and order harmonics, luni-solar perturbations, and solar radiation pressure. All of the perturbation effects were analyzed by secular, short, and long period variations for equinoctial orbit elements such as semimajor axis, eccentricity vector, inclination vector, and mean right ascension of the geostationary satellite. The reference stations for orbit determination were composed of four calibrated stations. Simulations were performed to evaluate the performance of real time analytical orbit determination in Korea. The simulation results demonstrated that it is possible to determine real time position of geostationary satellite with the accuracy of 300 m rms. This performance implies that the time accuracy is better than 25 ns all over the Korean peninsula. The real time analytical orbit determination method developed in this research can provide a reliable, extremely high accurate time synchronization service through setting up domestic-only benchmarks.

Orbit Determination Using Angle-Only Data for MEO & GEO Satellite and Obsolete

인공위성 궤도 결정을 위해 한국천문연구원의 0.6m 광시야 망원경을 이용하여 중 고궤도 인공위성과 폐기위성을 관측하였다. 관측 자료는 영상처리 및 좌표 보정을 통해 장착기기에 의한 오차를 보정한다. 위성 관측 사료에서 얻은 좌표 정보는 관측 시스템의 불안정성과 끝점 결정 오차에 의해 13각초의 오차를 가진다. KODAS의 결과로 얻은 시뮬레이션 좌표와 Gauss 방법을 이용해 예비궤도 결정을 수행하고 궤도 결정에 적합한 시간 간격을 찾아보았다. 또한 미분보정을 통한 예비궤도 결정 결과의 향상을 확인하였다. 이들 결과를 평균궤도요소 형대로 변환하여 실제 관측 자료와 비교하여 에비궤도 결정을 통해서 짧은 시간동안 위성의 추적이 가능함을 확인하였으며, 미분보정을 통해 그 결과를 향상시킬 수 있음을 확인하였다. 【We used an optical observation system with a 0.6m wide-field telescope and 5 computers system in KASI (Korean Astronomy and Space Science Institute) for satellite optical observation. Optical data have errors that are caused by targeting, expose start time and end-point determination. Gauss method for initial orbit determination was tested using angle-only data simulated by KODAS. And suitable time span is confirmed for result which has minimum errors. Initial orbit determination results are proved that optical observation system in KASI is possible satellite tracking for a short period. And also through differential correction, initial orbit determination results are improved.】

GEO Satellite Collision Avoidance Maneuver Strategy Against Inclined GSO Satellite

Orbit maneuver strategy for Geostationary Earth orbit (GEO) satellite is investigated to keep away from inclined geosynchronous orbit (GSO) satellite. Characteristics of inclined GSO with various combinations of eccentricity and argument of perigee are examined first. Then the close approach of inclined GSO SL-12 rocket body to GEO COMS satellite is inspected to develop a maneuver strategy for collision avoidance. Several sizes of deltavelocities are applied to the GEO satellite to check the effect of the maneuvers on separation. It is found that radial separation between the two satellites is the most important factor and the greatest separation can be achieved when the collision avoidance maneuver is executed at 12 hours before the time of closest approach.

Optical Orbit Determination of a Geosynchronous Earth Orbit Satellite Effected by Baseline Distances between Various Ground-based Tracking Stations II: COMS Case with Analysis of Actual Observation Data

We estimated the orbit of the Communication, Ocean and Meteorological Satellite (COMS), a Geostationary Earth Orbit (GEO) satellite, through data from actual optical observations using telescopes at the Sobaeksan Optical Astronomy Observatory (SOAO) of the Korea Astronomy and Space Science Institute (KASI), Optical Wide field Patrol (OWL) at KASI, and the Chungbuk National University Observatory (CNUO) from August 1, 2014, to January 13, 2015. The astrometric data of the satellite were extracted from the World Coordinate System (WCS) in the obtained images, and geometrically distorted errors were corrected. To handle the optically observed data, corrections were made for the observation time, light-travel time delay, shutter speed delay, and aberration. For final product, the sequential filter within the Orbit Determination Tool Kit (ODTK) was used for orbit estimation based on the results of optical observation. In addition, a comparative analysis was conducted between the precise orbit from the ephemeris of the COMS maintained by the satellite operator and the results of orbit estimation using optical observation. The orbits estimated in simulation agree with those estimated with actual optical observation data. The error in the results using optical observation data decreased with increasing number of observatories. Our results are useful for optimizing observation data for orbit estimation.

Alternating Sunspot Area and Hilbert Transform Analysis

We investigate the sunspot area data spanning from solar cycles 1 (March 1755) to 23 (December 2010) in time domain. For this purpose, we employ the Hilbert transform analysis method, which is used in the field of information theory. One of the most important advantages of this method is that it enables the simultaneous study of associations between the amplitude and the phase in various timescales. In this pilot study, we adopt the alternating sunspot area as a function of time, known as Bracewell transformation. We first calculate the instantaneous amplitude and the instantaneous phase. As a result, we confirm a ~22-year periodic behavior in the instantaneous amplitude. We also find that a behavior of the instantaneous amplitude with longer periodicities than the ~22-year periodicity can also be seen, though it is not as straightforward as the obvious ~22-year periodic behavior revealed by the method currently proposed. In addition to these, we note that the phase difference apparently correlates with the instantaneous amplitude. On the other hand, however, we cannot see any obvious association of the instantaneous frequency and the instantaneous amplitude. We conclude by briefly discussing the current status of development of an algorithm for the solar activity forecast based on the method presented, as this work is a part of that larger project.

Real time numerical dynamic orbit determination of geostationary satellite for time synchronization service

Abstract A real time numerical orbit determination method has been developed for a precise domestic time synchronization service. A one-way time synchronization method via a geostationary communication satellite was used. A differential method was also applied to mitigate errors. A numerical dynamic orbit determination method with a two-step estimator algorithm was implemented to further improve the differential mode time synchronization service through a more accurate acquisition of the geostationary satellite position. Perturbation force models were applied for the satellite dynamics, which include non-spherical geo-potential perturbations, lunar-solar perturbations, and solar radiation pressure. The reference stations for orbit determination were composed of four calibrated stations. Simulations were conducted to verify the performance of the proposed real time numerical orbit determination technique. The results verified that this technique made it possible to estimate real time position of a geostationary satellite within an accuracy of 2 km error in a spheroid radius. This performance demonstrates that the time accuracy is better than 40 nano seconds in the southern area of the Korean-peninsula. The real time numerical orbit determination method developed in this study can provide a reliable, highly accurate time transfer service through the establishment of domestic-only benchmarks.

Evolution of the Selenopotential Model and Its Effects on the Propagation Accuracy of Orbits around the Moon

The current work analyzes the effect of applying different selenopotential models to the propagation of a lunar orbiting spacecraft. A brief evolution history of the selenopotential model is first presented; then, four representative selenopotential models are selected for force modeling. Expected propagation errors are presented with respect to three different circular polar orbits around the Moon. As a result, an expected but rather significant number of orbit propagation errors are discovered. Compared to the solutions obtained using the GRAIL1500E model, the overall 3D propagation errors for a 4-day period could reach up to several tens of kilometers (50 km altitude case with the GLGM2 model) and up to several hundreds of meters (50, 100, and 200 km altitude cases even with the GRAIL660B model). For each different orbiter’s altitude, the appropriate ranges of the degree and order of the gravitational harmonic coefficients are also suggested to yield the best propagation performances with respect to the performance obtained with the full harmonic coefficients using the GRAIL1500E model. The results of the current work are expected to serve as practical guidelines for the field of system budget analysis, mission design, mission operations, and the analysis of scientific results.

Evaluating High-Degree-and-Order Gravitational Harmonics and its Application to the State Predictions of a Lunar Orbiting Satellite

In this work, an efficient method with which to evaluate the high-degree-and-order gravitational harmonics of the nonsphericity of a central body is described and applied to state predictions of a lunar orbiter. Unlike the work of Song et al. (2010), which used a conventional computation method to process gravitational harmonic coefficients, the current work adapted a well-known recursion formula that directly uses fully normalized associated Legendre functions to compute the acceleration due to the non-sphericity of the moon. With the formulated algorithms, the states of a lunar orbiting satellite are predicted and its performance is validated in comparisons with solutions obtained from STK/Astrogator. The predicted differences in the orbital states between STK/Astrogator and the current work all remain at a position of less than 1 m with velocity accuracy levels of less than 1 mm/s, even with different orbital inclinations. The effectiveness of the current algorithm, in terms of both the computation time and the degree of accuracy degradation, is also shown in comparisons with results obtained from earlier work. It is expected that the proposed algorithm can be used as a foundation for the development of an operational flight dynamics subsystem for future lunar exploration missions by Korea. It can also be used to analyze missions which require very close operations to the moon.

Preliminary Analysis of Delta-V Requirements for a Lunar CubeSat Impactor with Deployment Altitude Variations

Characteristics of delta-V requirements for deploying an impactor from a mother-ship at different orbital altitudes are analyzed in order to prepare for a future lunar CubeSat impactor mission. A mother-ship is assumed to be orbiting the moon with a circular orbit at a 90 deg inclination and having 50, 100, 150, 200 km altitudes. Critical design parameters that are directly related to the success of the impactor mission are also analyzed including deploy directions, CubeSat flight time, impact velocity, and associated impact angles. Based on derived delta-V requirements, required thruster burn time and fuel mass are analyzed by adapting four different miniaturized commercial onboard thrusters currently developed for CubeSat applications. As a result, CubeSat impact trajectories as well as thruster burn characteristics deployed at different orbital altitudes are found to satisfy the mission objectives. It is concluded that thrust burn time should considered as the more critical design parameter than the required fuel mass when deducing the onboard propulsion system requirements. Results provided through this work will be helpful in further detailed system definition and design activities for future lunar missions with a CubeSat-based payload.

Influence of the Choice of Lunar Gravity Model on Orbit Determination for Lunar Orbiters

We examine the influence of the lunar gravity model on the orbit determination (OD) of a lunar orbiter operating in a 100 km high, lunar polar orbit. Doppler and sequential range measurements by three Deep Space Network antennas and one Korea Deep Space Antenna were used. For measurement simulation and OD analysis, STK11 and ODTK6 were utilized. GLGM2, LP100K, LP150Q, GRAIL420A, and GRAIL660B were used for investigation of lunar gravity model selection effect. OD results were assessed by position and velocity uncertainties with error covariance and an external orbit comparison using simulated true orbit. The effect of the lunar gravity models on the long-term OD, degree and order level, measurement-acquisition condition, and lunar altitude was investigated. For efficiency verification, computational times for the five lunar gravity models were compared. Results showed that significant improvements to OD accuracy are observed by applying a GRAIL-based model; however, applying a full order and degree gravity modeling is not always the best strategy, owing to the computational burden. Consequently, we consider that OD using GRAIL660B with 70 × 70 degree and order is the most efficient strategy for mission preanalysis. This study provides useful guideline for KPLO OD analysis during nominal mission operation.