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Abstract: Autonomous and magnetometer-based orbit determination algorithm in target pointing satellite based solely on magnetometer measurements for speeding up filter convergence is proposed. The cosine angle between geomagnetic field vector and target direction vector, as viewed from the satellite, provides some quantity that is sensitive to satellite position, more regular and periodical than the magnetic field magnitude as a satellite moves around its orbit, and easily observable from magnetometers. An algorithm is formulated by applying the unscented Kalman filter (UKF) based on the laws of orbital mechanics, the measurement models of geomagnetic field magnitude and the cosine angle. The algorithm is tested and verified by using real-flight data obtained from Korea Multi-Purpose Satellite-1 (KOMPSAT-1) and Magnetic Field Satellite (MAGSAT). The results obtained from numerical simulations using KOMPSAT-1 and MAGSAT data illustrate the significant convergence speed up and the robustness against both noise and bias of magnetic measurements. The proposed orbit determination is adequate as a primary routine for a target pointing satellite mission that requires fast convergence. Otherwise, it is sufficient as a secondary routine to provide redundancy. [Copyright &y& Elsevier]

• The main contribution of this study is the development of a new sequential-convex-programming-based technique that can periodically update the real-time guidance solution in consideration of the minimum-fuel Mars powered descent guidance problem with free final time. • To illustrate the validity of the proposed strategy for on-board real-time applications, an on-board test was conducted using the GR740, which is the processor of choice for the next-generation on-board computers of the European Space Agency (ESA) and Korea Aerospace Research Institute (KARI). • To assess the performance of the strategy, numerical simulations were conducted under various conditions. • Furthermore, the results of Monte Carlo simulations revealed that the proposed strategy is adequately robust against environmental and systematic disturbances. An on-board guidance generation approach for Mars pinpoint landing has been developed based on sequential convex programming. To optimize the flight time in the formulation, a discretized form of the landing problem was constructed under the conditions of linearized states, controls, and time increment between consecutive time steps. To implement this strategy in real time, on-board demonstrations were conducted with the GR740, which is the next-generation on-board processor of choice for the European Space Agency. The total number of time steps was determined based on the results of these demonstrations. The numerically simulated results also indicate that the solution obtained via the proposed strategy is close to the optimized GPOPS-II solution under the condition of no disturbance. Even under disturbances such as navigation errors, initial prediction errors, and perturbation forces, the proposed strategy ensures that the spacecraft reaches its target position with near-optimal fuel consumption. [ABSTRACT FROM AUTHOR]

Satellite dc–dc power converters are required to have electromagnetic interference (EMI) compatibility as well as high reliability. The requirement of high reliability limits the complexity of the structure as well as the number of low-reliability components. In this paper, a conventional active-clamp-forward (ACF) topology is examined for a satellite converter owing to its relatively simple structure, small number of components, and good clamping capability regarding its switch voltage stress. However, it has a pulsating input current shape, a high di/dt level, and high switch voltage stress, which are disadvantageous with regard to the EMI filter size and high-input-voltage converter applications. To overcome these drawbacks, a new ACF topology with a continuous input current waveform is proposed. In the proposed waveform, the voltage stresses on the main switches are relieved, which is crucial to ensure the reliability of satellite FET switches, by utilizing a two-series connected structure. These features allow the proposed converter to serve as a high-input-voltage, high power-density satellite converter. In this paper, 100 W prototype converters with 50 V inputs operating at 300 kHz are built with a high power density of 47.9 W/in3. Experiments on these prototype converters are carried out to verify the features. [ABSTRACT FROM AUTHOR]

Abstract A Control Moment Gyro (CMG) cluster for an agile spacecraft usually suffers from a singularity problem where it does not meet a desired torque. To solve the problem, a proposed steering law consists of SR (singularity robustness) inverse and SR null motion to compensate each other. The SR null motion solves the singularity problem that the SR inversestays in the singularities. Especially, an APF (artificial potential function) is included in the SR null motion to avoid internal singularities and to reach external singularity simultaneously. The APF newly considers the distance between the current gimbal angle and the gimbal angle corresponding to all singularities. All the combinations of singular gimbal angles are regarded as a target and obstacles of the potential function. After a spacecraft maneuver, an initial gimbal angle is regarded as the target to secure the repeatability of next maneuver. When avoiding or escaping the obstacles corresponding to the unwanted singular gimbal angles, the proposed steering law rapidly changes the gimbal angle by taking torque error. To overcome the local minima problem which is generally a serious drawback in using the potential function, an algorithm for changing parameters of the potential function are proposed in the steering law. To avoid the internal singularity, moreover, the present study introduces a configuration change of the CMG cluster. The proposed steering law is mathematicallyanalyzed and verified through numerical simulations. It shows that the CMG cluster can avoid an internal singularity rapidly, resulting in faster maneuver time of the spacecraft. Highlights • A novel singularity avoidance steering law of control moment gyros. • Artificial potential function (APF) multiplied by singularity robustness null motion. • Consider singular gimbal angles or initial gimbal angles as obstacles or a goal of APF. • Solution to solve local minima problem and to utilize modified pyramid configuration. • Mathematical and simulation analysis of the proposed steering law. [ABSTRACT FROM AUTHOR]