Kyu-Hong Choi

Nonlinear Attitude Control of a Tether-Connected Multi-Satellite in Three-Dimensional Space

The objective of the current research is to analyze the attitude dynamics and control of a tethered satellite formation flying where the tethered units are modeled as extended rigid bodies. The three-inline array system is used in this study, and the general formulation of the equations of motion of the system is obtained through a Lagrangian approach. In this research, attitude motions of the tethered satellite system are analyzed in a three-dimensional free-space system to complement previous works. The state-dependent Riccati equation (SDRE) controller is used to regulate the attitude errors. The stability region for the SDRE-controlled tethered satellite system is also estimated using a numerical method to show globally asymptotic stability for the control method. Centralized and decentralized approaches are applied to the dynamic system to compare the performance of controlling the attitude motion. The SDRE controller performs well in both the centralized and decentralized approaches for the attitude control of tethered satellites in formation flying.

Spacecraft orbit determination using GPS navigation solutions

Abstract The orbit determination using the GPS navigation solutions for the KOMPSAT-1 spacecraft has been studied. The Cowell method of special perturbation theories was employed to develop a precision orbit propagation, and the perturbations due to geopotential, the gravity of the Sun and the Moon, solid Earth tides, ocean tides, the Earths dynamic polar motion, solar radiation pressure, and atmospheric drag were modeled. Specifically, the satellite box-wing macro model was applied to minimize the drag errors at low altitude. The estimation scheme consisted of an extended Kalman filter and Bayesian least square method. To investigate the applicability of the method to the KOMPSAT-1 spacecraft, the orbit determination was accomplished using the GPS navigation solutions for the TOPEX/POSEIDON and TAOS satellites. The orbit determination results were compared with NASA POE generated by global laser tracking. The position and velocity accuracy was estimated about 16∼7 m and 0.0157∼0.0074 m·s−1 RMS, respectively, for the two satellites in the presence of SA. These results verify that an orbit determination scheme using GPS navigation solutions can provide the static orbit information and reduce conspicuously the position and velocity errors of navigation solutions. It can be suggested that the sequential and batch orbit determination using the GPS navigation solutions be the most appropriate method in the KOMPSAT-1 type mission.

Orbit Determination Using the Geomagnetic Field Measurement via the Unscented Kalman Filter

OrbitdeterminationofspacecraftusingonlymagnetometermeasurementsviatheunscentedKalman filter(UKF) is presented. An algorithm was formulated by adopting the UKF and an adequate dynamic model developed for processing geomagnetic field measurements. The paper consists of the analysis of force and estimation models, the dependencyofposition accuracyonorbittype,andmeasurement errors,as wellas acomparisonof theUKFandthe extended Kalman filter (EKF). Finally, the developed algorithm used actual magnetometer flight data from the Magnetic Field Satellite (MAGSAT). The results obtained from the MAGSAT data demonstrate that the achieved position error was approximately 2 km. The UKF performs similar to the EKF in position accuracy for a sampling interval of less than 20 s. In contrast, for a sampling interval of 40 s, the EKF yields lower position accuracy than the UKF.Inparticular,thedecomposedpositionerrorsarenotbiasedbecausetheUKFisnotaffectedbylinearizationof the measurement function. This improvement makes the magnetometer-based orbit determination method more robust and reliable as a real-time orbit determination system for small satellites that require moderate position accuracy, and also as a backup orbit determination system for large satellites. Nomenclature adrag = acceleration due to atmospheric drag, m=s 2 ageo = acceleration due to geopotential, m=s 2 aSRP = acceleration due to the solar radiation pressure, m=s 2 asun=moon = acceleration due to the sun and moon’s gravity, m=s 2 Br, B� , B� = the Earth magnetic field vector components, nT B � = inverse of the ballistic coefficient, m 2 =kg e = eccentricity

Optimal low-thrust intercept/rendezvous trajectories to earth-crossing objects

Optimal trajectories are presented that intercept and rendezvous with Earth-crossing objects by using an advanced magnetoplasma spacecraft with variable low-thrust capability. A detailed optimization method is formulated to provide optimal trajectories for any threat scenarios caused by Earth-crossing asteroids/comets. The characteristics of the trajectories are analyzed to find insights into appropriate trajectories for a targeted celestial body. It is primarily illustrated that the optimal trajectories are significantly dependent on the orbital elements of the impacting object, as well as on the distance from the Earth to the target. The trajectories can be used as reference trajectories to formulate an overall concept for solving the deflection problem of the Earth-crossing objects.

Analysis of a station-keeping maneuver strategy for collocation of three geostationary satellites

Abstract A collocation strategy has been planned and analyzed for three geostationary satellites in the same longitude control box. The orbit determination error analysis is performed when only one ground station is used for the angle tracking and ranging. Based on the orbit determination errors, the station-keeping bands are allocated for seven-day East/West and 14-day North/South station-keeping maneuver cycles. The eccentricity control circle and inclination control box for individual satellites are allocated for the collocation. The eccentricity vector and inclination vector separation method is applied for the collocation, and the maneuver schedule is planned to minimize the operational load by avoiding simultaneous maneuvers. A total of fourteen weeks of station-keeping maneuvers are performed for three KOREASAT satellites, collocated in 116°E±0.05° longitude band.

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 of Spacecraft Using Global Positioning System Single-Frequency Measurement

The dynamic orbit determination of a low Earth orbiter using global positioning system single-frequency measurements has been implemented. Currently two methods are being applied to eliminate or reduce ionospheric path delay in single-frequency measurement. One is a group and phase ionosphere calibration technique using code pseudorange and L1 carrier phase, and the other is application of total electron content values from an ionospheric model using only L1 carrier phase to determine the orbit. A new method based on the latter has been developed, which estimates the scale factors of total electron content values in the location of a low Earth orbiter once per each measurement time. Orbit determination using actual global positioning system measurements of the TOPEX/POSEIDON and the Challenging minisatellite payload was conducted to verify the accuracy of the new method. It is verified that, if the total electron contents scale factor estimation technique were applied, 1-m level position accuracy (1σ) for low Earth orbit below 500-km altitudes could be achieved using precision orbit determination based on the global positioning system double-differencing method.

Substorms associated with azimuthal turnings of the interplanetary magnetic field

Abstract Whether the magnetospheric substorms can be triggered by the interplanetary magnetic field (IMF) variations is an important issue in the substorm research. In this work we investigate observationally the relationship between substorm activities and IMF By variations, i.e., azimuthal turnings. We have searched for the IMFs azimuthal turning events for a period of one year using the data from multispacecraft monitoring the solar wind, WIND, IMP 8 and Geotail. Based on specific selection criteria, we have found 11 such events that exhibit pure azimuthal turnings while the IMF Bz remains quasi-steady. These events are found to be mostly in reasonable temporal associations with the substorm activities which were identified by multipoint measurements using the geomagnetic bays, auroral images, geosynchronous energetic particle injections and magnetic dipolarizations. We find an average response time of ∼8– 9 min between the substorm onsets and the dayside magnetopause contact times of the IMF By turnings. The results suggest that in addition to the more popular trigger by northward turnings of the IMF, its azimuthal (both positive and negative) turnings be also regarded as another possible external trigger of substorms.

Collocation of two GEO satellites and one inclined GSO satellite

Abstract Three collocation strategies are planned and analyzed for the cluster of two geostationary orbit (GEO) satellites and one inclined geosynchronous orbit (GSO) satellite in the same longitude control band of 116°E±0.05° . The longitudinal control bands are allocated for the two GEO satellites and one inclined GSO satellite with seven-day East/West station-keeping maneuver cycle. The latitudinal control bands are allocated for the two GEO satellites with fourteen-day North/South station-keeping maneuver cycle. One inclined GSO satellite is allowed for natural inclination drift. The coordinated eccentricity vector and inclination vector separation method is applied for the collocation, and the maneuver schedule is planned to minimize the operational load by avoiding simultaneous maneuvers. A total of six months of station-keeping maneuver simulations are performed for the three different strategies.

Orbital rendezvous using two-step sliding mode control

Abstract The problem of spacecraft rendezvous is studied, using sliding mode control in the presence of the earths gravitational perturbation. The impulsive solution of Lamberts problem is obtained using the combined equations method to minimize total ΔV via an iterative method, and it is used as the desired trajectory for the rendezvous. In this paper, a two-step sliding mode control method is introduced for solving the rendezvous problem with finite-thrust including unmodeled dynamics. The thrust-coast-thrust type control laws for the system to follow the desired trajectories are represented and resultant trajectories are close enough to the Lamberts orbit with comparable amount of ΔV to the Lamberts impulsive solution. All state variables matched the final boundary conditions reasonably well at the end of maneuver.