A new approach to the telescope operation method for satellite tracking using a time synchronization technique

For the development of a telescope that is capable of precisely tracking satellites and high-speed operation such as satellite laser ranging, a special method of telescope operation is required. This study aims to propose a new telescope operation method and system configuration for the independent development of a mount and an operation system which includes the host computer. Considering that the tracking of a satellite is performed in real time, communication and synchronization between the two independent subsystems are important. Therefore, this study applied the concept of time synchronization, which is used in various fields of industry, to the communication between the command computer and the mount. In this case, communication delays do not need to be considered in general, and it is possible to cope with data loss. Above all, when the mount is replaced in the future, only the general communication interface needs to be modified, and thus, it is not limited by replacement in terms of the overall system management. The performance of the telescope operation method developed in this study was verified by applying the method to the first mobile SLR system in Korea. This study is significant in that it proposed a new operation method and system configuration, to which the concept of time synchronization was applied, for the observation system that requires an optical telescope.     

The Scientific Measurement System of the Gravity Recovery and Interior Laboratory (GRAIL) Mission

The Gravity Recovery and Interior Laboratory (GRAIL) mission to the Moon utilized an integrated scientific measurement system comprised of flight, ground, mission, and data system elements in order to meet the end-to-end performance required to achieve its scientific objectives. Modeling and simulation efforts were carried out early in the mission that influenced and optimized the design, implementation, and testing of these elements. Because the two prime scientific observables, range between the two spacecraft and range rates between each spacecraft and ground stations, can be affected by the performance of any element of the mission, we treated every element as part of an extended science instrument, a science system. All simulations and modeling took into account the design and configuration of each element to compute the expected performance and error budgets. In the process, scientific requirements were converted to engineering specifications that became the primary drivers for development and testing. Extensive simulations demonstrated that the scientific objectives could in most cases be met with significant margin. Errors are grouped into dynamic or kinematic sources and the largest source of non-gravitational error comes from spacecraft thermal radiation. With all error models included, the baseline solution shows that estimation of the lunar gravity field is robust against both dynamic and kinematic errors and a nominal field of degree 300 or better could be achieved according to the scaled Kaula rule for the Moon. The core signature is more sensitive to modeling errors and can be recovered with a small margin.     

Nanosatellite constellation deployment using on-board magnetic torquer interaction with space plasma

One of the advantages that drive nanosatellite development is the potential of multi-point observation through constellation operation. However, constellation deployment of nanosatellites has been a challenge, as thruster operations for orbit maneuver were limited due to mass, volume, and power. Recently, a de-orbiting mechanism using magnetic torquer interaction with space plasma has been introduced, so-called plasma drag. As no additional hardware nor propellant is required, plasma drag has the potential in being used as constellation deployment method. In this research, a novel constellation deployment method using plasma drag is proposed. Orbit decay rate of the satellites in a constellation is controlled using plasma drag in order to achieve a desired phase angle and phase angle rate. A simplified 1D problem is formulated for an elementary analysis of the constellation deployment time. Numerical simulations are further performed for analytical analysis assessment and sensitivity analysis. Analytical analysis and numerical simulation results both agree that the constellation deployment time is proportional to the inverse square root of magnetic moment, the square root of desired phase angle and the square root of satellite mass. CubeSats ranging from 1 to 3?U (1–3?kg nanosatellites) are examined in order to investigate the feasibility of plasma drag constellation on nanosatellite systems. The feasibility analysis results show that plasma drag constellation is feasible on CubeSats, which open up the possibility of CubeSat constellation missions.     

Updated TUS space fluorescence detector for study of UHECR

The TUS (Tracking Ultra-violet Set up) space fluorescence detector has to be launched in 2010 as a separated platform in Foton (Bion) mission prepared by the Samara enterprise. This detector was designed for another satellite and the updated variant of the TUS detector for a new platform is presented. The data on UV glow of the atmosphere obtained in operation of one pixel of the TUS detector on board the Moscow State University “Universitetsky-Tatiana” satellite was taken into account in design of the updated TUS detector. The data on UV transient flashes registered in “Universitetsky-Tatiana” mission are of special interest. Electronics of the TUS detector able to select and register different types of UV events in the atmosphere is presented.     

A Bound on Processor Gain

In data transmission systems using electromagnetic propagation there is often interest in the antijam capability of the system and the probability that the transmitted signal can be intercepted. It is shown that, neglecting antenna gain, both of these are adequately described by the processor gain. An information theoretic bound is given for the processor gain and the performance of several conventional signaling systems is indicated. Improvement using forward error correction is also investigated.     

The Dawn Gravity Investigation at Vesta and Ceres

The objective of the Dawn gravity investigation is to use high precision X-band Doppler tracking and landmark tracking from optical images to measure the gravity fields of Vesta and Ceres to a half-wavelength surface resolution better than 90-km and 300-km, respectively. Depending on the Doppler tracking assumptions, the gravity field will be determined to somewhere between harmonic degrees 15 and 25 for Vesta and about degree 10 for Ceres. The gravity fields together with shape models determined from Dawn??s framing camera constrain models of the interior from the core to the crust. The gravity field is determined jointly with the spin pole location. The second degree harmonics together with assumptions on obliquity or hydrostatic equilibrium may determine the moments of inertia.     

Hardware-in-the-loop simulations of GPS-based navigation and control for satellite formation flying

A relative navigation and formation control algorithm for satellite formation flying was developed, and a hardware-in-the-loop (HIL) simulation testbed was established and configured to evaluate this algorithm. The algorithm presented is a relative navigation estimation algorithm using double-difference carrier-phase and single-difference code measurements based on the extended Kalman filter (EKF). In addition, a state-dependent Riccati equation (SDRE) technique is utilized as a nonlinear controller for the formation control problem. The state-dependent coefficient (SDC) form is formulated to include nonlinearities in the relative dynamics. To evaluate the relative navigation and control algorithms developed, a closed-loop HIL testbed is configured. To demonstrate the performance of the testbed, a test formation flying scenario comprising formation acquisition and keeping in a low earth orbit (LEO) has been established. The relative navigation results from the closed-loop simulations show that a 3D RMS of 0.07 m can be achieved for position accuracy. The targeted leader–follower formation flying in the along-track separation of 100 m was maintained with a mean position error of approximately 0.2 m and a standard deviation of 0.9 m. The simulation results show that the HIL testbed is capable of successful demonstration of the GPS-based satellite autonomous formation flying mission.     

Buried small objects detected by UWB GPR

A ground penetrating radar (GPR) using short-pulse is developed to detect small and shallow metal objects buried underground. A bistatic mode in which the GPR system uses separate transmitting and receiving antennas is applied. A modified fat dipole antenna is developed for the transmitting and receiving antennas. The prototype of the system is tested in the real environment and 2D visualization of raw data is achieved. We show that the developed system has a good ability to detect underground metal objects, and even small targets of several centimeters.     

Investigating suitable orbits for the Swarm constellation mission – The frozen orbit

This paper proposes a suitable orbit design for the lower pair of ESA's Swarm constellation mission, flying side-by-side in near-polar and circular orbits with a separation of only 1.4° at ascending node. Both orbits are suggested to be frozen orbits to minimize the evolution, and an along-track separation strategy is applied to avoid collision risk. The characteristics of the proposed orbit type are examined through numerical techniques including high-fidelity perturbation models. The prime change from the initial configuration is an along-track separation. The perturbations causing the along-track drift are analyzed by switching on/off certain perturbations. The results indicate that the tesseral harmonics and the atmospheric drag yield dominant effects. The atmospheric drag effect shows a dependence on the local time of the ascending node. From two months of orbit propagation for the altitude 300 km the maximum along-track drift we obtain is about 80 km, which is still within the measurement requirement range. Several maneuver strategies for maintaining the proposed orbit design are suggested. The results analyzed for the proposed orbit design show that collision risk can be avoided by along-track separation within the frozen orbit design. Consequently, this combination is considered as a suitable approach for Swarm's lower pair.