Browsing by Author "Mughal, Muhammad Rizwan"

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    Coulomb drag propulsion experiments of ESTCube-2 and FORESAIL-1
    (Elsevier Limited, 2020-12) Iakubivskyi, Iaroslav; Janhunen, Pekka; Praks, Jaan; Allik, Viljo; Bussov, Kadri; Clayhills, Bruce; Dalbins, Janis; Eenmäe, Tõnis; Ehrpais, Hendrik; Envall, Jouni; Haslam, Sean; Ilbis, Erik; Jovanovic, Nemanja; Kilpua, Emilia; Kivastik, Joosep; Laks, Jürgen; Laufer, Philipp; Merisalu, Maido; Meskanen, Matias; Märk, Robert; Nath, Ankit; Niemelä, Petri; Noorma, Mart; Mughal, Muhammad Rizwan; Nyman, Samuli; Pajusalu, Mihkel; Palmroth, Minna; Paul, Aditya Savio; Peltola, Tatu; Plans, Mathias; Polkko, Jouni; Islam, Quazi Saimoon; Reinart, Anu; Riwanto, Bagus; Sammelselg, Väino; Sate, Janis; Sünter, Indrek; Tajmar, Martin; Tanskanen, Eija; Teras, Hans; Toivanen, Petri; Vainio, Rami; Väänänen, Mika; Slavinskis, Andris; Department of Electronics and Nanoengineering; Jaan Praks Group; Eija Tanskanen Group; Esa Kallio Group; University of Tartu; Finnish Meteorological Institute; Estonian Student Satellite Foundation - ESTCube; Department of Electronics and Nanoengineering; University of Helsinki; Technical University of Dresden; Jaan Praks Group; Ventspils University College; University of Turku
    This paper presents two technology experiments – the plasma brake for deorbiting and the electric solar wind sail for interplanetary propulsion – on board the ESTCube-2 and FORESAIL-1 satellites. Since both technologies employ the Coulomb interaction between a charged tether and a plasma flow, they are commonly referred to as Coulomb drag propulsion. The plasma brake operates in the ionosphere, where a negatively charged tether deorbits a satellite. The electric sail operates in the solar wind, where a positively charged tether propels a spacecraft, while an electron emitter removes trapped electrons. Both satellites will be launched in low Earth orbit carrying nearly identical Coulomb drag propulsion experiments, with the main difference being that ESTCube-2 has an electron emitter and it can operate in the positive mode. While solar-wind sailing is not possible in low Earth orbit, ESTCube-2 will space-qualify the components necessary for future electric sail experiments in its authentic environment. The plasma brake can be used on a range of satellite mass classes and orbits. On nanosatellites, the plasma brake is an enabler of deorbiting – a 300-m-long tether fits within half a cubesat unit, and, when charged with -1 kV, can deorbit a 4.5-kg satellite from between a 700- and 500-km altitude in approximately 9–13 months. This paper provides the design and detailed analysis of low-Earth-orbit experiments, as well as the overall mission design of ESTCube-2 and FORESAIL-1.
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    A Detailed Thermal and Effective Induced Residual Spin Rate Analysis for LEO Small Satellites
    (IEEE-INST ELECTRICAL ELECTRONICS ENGINEERS INC, 2020-01-01) Ali, Anwar; Tong, Jijun; Ali, Haider; Mughal, Muhammad Rizwan; Reyneri, Leonardo M.; Department of Electronics and Nanoengineering; Jaan Praks Group; Zhejiang Sci-Tech University; University of Technology, Nowshera; Polytechnic University of Turin
    In space thermal environment, satellites are exposed to multiple heat sources which can deteriorate structural and equipment integrity over long periods of time. Normally radiators are used to release heat, but due to space and weight constraints, it is impossible to mount radiators on small satellites. This problem signifies the importance of thermal analysis of a satellite in every development stage, such as design, manufacturing and testing. The ultimate goal of this work is to analyze a small spacecraft in space thermal environment by considering the effect of various heat sources. Thermal equilibrium equation is achieved which is applied to spacecraft with different shapes and dimensions and temperature is measured for a range of absorption co-efficient values (i.e. 0.5 similar to 0.9). Through an experimental setup a method is devised to measure the absorption co-efficient of small satellites that can be used for exact temperature measurement. Secondly, the paper presents a preliminary analysis of induced spin produced by small satellites due to asymmetrical colors (different absorptance) of satellite outer surface. The substantial contributors for induced spin are considered and the estimated spin is measured.
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    Particle swarm optimization for magnetometer calibration with rotation axis fitting using in-orbit data
    (IEEE-INST ELECTRICAL ELECTRONICS ENGINEERS INC, 2022-04) Riwanto, Bagus Adiwiluhung; Niemela, Petri; Ehrpais, Hendrik; Slavinskis, Andris; Mughal, Muhammad Rizwan; Praks, Jaan; Department of Electronics and Nanoengineering; Jaan Praks Group; University of Tartu
    This article demonstrates the performance of an improved particle swarm optimization (PSO) algorithm with scalar checking and rotation axis fitting objectives using in-orbit data, which is obtained from two CubeSats missions, Aalto-1 and ESTCube-1, as well as simulation as reference. The improved algorithm uses sequential objectives refinement process to combine the two optimization objectives. This improvement addresses some challenges of magnetometer calibration when using in-orbit data. First, the change in the magnetic field vector direction at different points in orbit which is uncorrelated to the rotation of the spacecraft itself. Second, the uncertainty of the rotation axis information used as the reference, e.g., from gyroscope noise. Third, the available data set is heavily affected by the rotation mode of the spacecraft, which imposes some limitation in the rotation axis information needed by the algorithm. The improved PSO algorithm is applied on simulated data in order to analyze the calibration performance under different spacecraft tumbling rates and noise levels. In ideal condition (varying rotation axis during measurements and sufficient sampling rate relative to the spin rate), the rotation axis fitting objective can reach ∼0.1° of correction accuracy.
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    Verification of Tether Deployment System aboard CubeSat through Dynamics Simulations and Tests
    (2021-03-06) Sakamoto, Hiraku; Mughal, Muhammad Rizwan; Slavinskis, Andris; Praks, Jaan; Toivanen, Petri; Janhunen, Pekka; Palmroth, Minna; Kilpua, Emilia; Vainio, Rami; Department of Electronics and Nanoengineering; Jaan Praks Group; Finnish Meteorological Institute; University of Helsinki; University of Turku; Tokyo Institute of Technology
    This paper proposes a proper model selection strategies for the dynamic simulations of the tether deployment mission aboard a CubeSat. Space tether technology will enable innovative space missions in the near future. The Coulomb Drag Propulsion (CDP), including electric solar wind sailing, is one of the plausible future technologies. The authors currently develop a CubeSat, FORESAIL-1, for space demonstration of CDP. However, the analytical simulations for the verification and validation of the mission design typically require a high computational cost. This is because a minimum model order is not selected properly. In this study, through observing a preliminary analytical model for tether deployment analysis, the simplest model is chosen to avoid the mission failure modes in each deployment phase.