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Journal articles on the topic 'Magnetorquer'

Author: Grafiati

Published: 4 June 2021

Last updated: 20 February 2023

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1

Bai, Bo, Jun Zhou, and Shengyun Wang. "Design of High-Performance Magnetorquer with Air Core for CubeSat." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 36, no. 1 (February 2018): 1–6. http://dx.doi.org/10.1051/jnwpu/20183610001.

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To solve the problem that how to design a big magnetic moment, small size, light weight, low power consumption magnetorquer with air core under the constraint of limited volume and power in CubSate, multiobjective optimization design method is used. Firstly, based on the structure of the square support with multiple layer of the enameled wire wrapped, the magnetic moment model, power consumption model and mass model are deduced from square support size, enameled wire diameter and turn number, respectively. Secondly, according to the model of magnetic moment and power consumption, the multi-objective optimization design of magnetorquer with genetic algorithm (GA) is used under the constraint of limited mass and volume. Thirdly, based on the relation between the magnetic moment and the magnetic induction intensity, the measurement method of the magnetic moment is designed. Finally, the designed parameter is implemented. The test result showed that the designed magnetorquer has the qualities of high linearity, low remanence, and it well met the requirements of CubeSat standard. The designed magnetorquer successfully applied in several CubeSats indicates the reliability of this design scheme.

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2

Si, Juntian, Yang Gao, and Abadi Chanik. "Slew Control of Prolate Spinners Using Single Magnetorquer." Journal of Guidance, Control, and Dynamics 39, no. 3 (March 2016): 719–27. http://dx.doi.org/10.2514/1.g001035.

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3

MIYATA, Kikuko, Tomohiro NARUMI, and Jozef C. van der HA. "Comparison of Different Magnetorquer Control Laws for QSAT." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, SPACE TECHNOLOGY JAPAN 7, ists26 (2009): Pd_43—Pd_48. http://dx.doi.org/10.2322/tstj.7.pd_43.

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4

Kuiper, Hans, and Dennis Dolkens. "A cutting edge 6U CubeSat ADCS design for Earth observation with sub-meter spatial resolution at 230–380 km altitude." CEAS Space Journal 12, no. 4 (June 18, 2020): 613–21. http://dx.doi.org/10.1007/s12567-020-00323-7.

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Abstract A 6U CubeSat for Earth observation in 230–350 km orbits with sub-meter resolution is presented. The proposed Stable and Highly Accurate Pointing Earth-Imager (SHAPE) system’s attitude determination and control system (ADCS) is composed of a single momentum bias wheel with magnetic bearings at rotational speeds of 6000–7000 rpm and refined magnetorquers. Reaction wheels as instability source are absent. The ADCS stabilizes the spacecraft attitude by counteracting the torques from external disturbances in the thermosphere down to < 1° pointing accuracy and < 0.1° instability. The momentum wheel was sized to an angular momentum of 1 Nms based on the worst-case atmospheric density of the next solar cycle. The 0.5 Am2 magnetorquer dipole moment provides with low power consumption, mass and cost, high reliability and sufficient torque. The ADCS initialisation study revealed three stable start-up modes, while the all-spun state is achieved using a set of thrusters. De-tumbling analysis show that the magnetorquers reduce the tumbling rates with magnitudes of up to 35°/s to mean motion values in less than an orbit using a static gain B-dot controller. A 3U camera design capable of sub-meter spatial resolution at 230 km altitude is presented which complies with the SHAPE spacecraft system design. The instrument has a single deployable primary mirror enabled by a deployment hinge design with hysteresis < 0.5 μ. This payload combined with air-breathing electric propulsion technology at 230 km nominal altitude boosts the SHAPE system Earth observation potential down to sub-meter spatial resolution and enables tuning of the mission lifetime by orbit keeping.

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5

Hou, Xu Guang, Jian Yan, Jin Jin, and Shun Liang Mei. "Magnetorquer Based Vertical Damping Method for Microsatellite Attitude Control." Applied Mechanics and Materials 263-266 (December 2012): 584–87. http://dx.doi.org/10.4028/www.scientific.net/amm.263-266.584.

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Aiming at a three-axis stabilized microsatellite, a novel attitude control method, called magnetorquer based vertical damping, is proposed to avoid the occurrence of the worst situation that the non-solar-battery-plane spins towards the sun. DSP based simulation results based on DSP show that the vertical damping method outperforms the simple damping method when no orbit information is available, simultaneously the whole attitude control scheme is simple and effective. The proposed solution guarantees a stable power supply from the electrical source even under the extreme situation, which improves the reliability of the whole microsatellite system.

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6

Ali, Anwar, Shoaib Ahmed Khan, M. Usman Khan, Haider Ali, M. Rizwan Mughal, and Jaan Praks. "Design of Modular Power Management and Attitude Control Subsystems for a Microsatellite." International Journal of Aerospace Engineering 2018 (December 17, 2018): 1–13. http://dx.doi.org/10.1155/2018/2515036.

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The Electric Power System (EPS) and attitude control system (ACS) are the essential components of any satellite. EPS and ACS efficiency and compactness are substantial for the proper operation and performance of the satellite’s entire mission life. So, realizing the significance of EPS and ACS subsystems for any satellite, they have been assimilated and developed in modular forms focusing on efficiency and compactness. The EPS is comprised of three modules called the solar panel module (SPM), power conditioning module (PCM), and power distribution module (PDM) while the ACS has an embedded magnetorquer coil. For compactness and miniaturization purposes, the magnetorquer coil is embedded inside the SPM. The components used are commercial off-the-shelf (COTS) components emphasizing on their power efficiency, small dimensions, and weight. Latch-up protection systems have been designed and analyzed for CMOS-based COTS components, in order to make them suitable for space radioactive environment. The main design features are modularity, redundancy, power efficiency, and to avoid single component failure. The modular development of the EPS and ACS helps to reuse them for future missions, and as a result, the overall budget, development, and testing time and cost are reduced. A specific satellite mission can be achieved by reassembling the required subsystems.

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7

Utama, S., P. R. Hakim, and M. Mukhayadi. "Quarter orbit maneuver using magnetorquer to maintain spacecraft angular momentum." IOP Conference Series: Earth and Environmental Science 284 (May 31, 2019): 012046. http://dx.doi.org/10.1088/1755-1315/284/1/012046.

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8

Kondo, Kota, Ilya Kolmanovsky, Yasuhiro Yoshimura, Mai Bando, Shuji Nagasaki, and Toshiya Hanada. "Nonlinear Model Predictive Detumbling of Small Satellites with a Single-Axis Magnetorquer." Journal of Guidance, Control, and Dynamics 44, no. 6 (June 2021): 1211–18. http://dx.doi.org/10.2514/1.g005877.

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9

Ali, Anwar, M. Rizwan Mughal, Haider Ali, Leonardo M. Reyneri, and M. Naveed Aman. "Design, implementation, and thermal modeling of embedded reconfigurable magnetorquer system for nanosatellites." IEEE Transactions on Aerospace and Electronic Systems 51, no. 4 (October 2015): 2669–79. http://dx.doi.org/10.1109/taes.2015.130621.

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10

Asadabadi, Amirhossein, and Amir M. Anvar. "Small Satellite Modelling and Three-Axis Magnetorquer-Based Stabilisation Using Fuzzy Logic Control." Applied Mechanics and Materials 152-154 (January 2012): 1639–44. http://dx.doi.org/10.4028/www.scientific.net/amm.152-154.1639.

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Recently small satellites have become increasingly popular because of their ability to provide educational institutes with the chance to design, construct, and test their spacecraft from beginning to the possible launch due to the low launching cost and development of microelectronics (Figure 1). Clearly, using only electromagnetic coils instead of different types of actuators will serve the purpose of weight reduction where every grams counts. But some restrictions described in the paper limit utilising only “Electromagnetic” actuation for 3D stabilisation and adversely affects the efficiency of the controller. However, there are some theories developed recently that have made the aforementioned purpose feasible. In this paper a new control method based on Fuzzy Logic Control (FLC) is presented that keeps the satellite in desired conditions only by electromagnetic coils. More precisely, an approach of Fuzzy control which is incorporated with electromagnetic actuation is presented for the in-orbit attitude control of a small satellite. The design is developed to stabilize the spacecraft against disturbances with a three-axis stabilizing capability. The paper also describes the required hardware and the design and development of the magnetic torquers.

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11

Ahmed Khan, Shoaib, Anwar Ali, Yang Shiyou, Shah Fahad, and Jijun Tong. "Optimized Design and Analysis of Printed Magnetorquer for a 3-U Nano-Satellite." Journal of Aerospace Engineering 35, no. 1 (January 2022): 04021103. http://dx.doi.org/10.1061/(asce)as.1943-5525.0001343.

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12

Mughal, Muhammad Rizwan, Hassan Ali, Anwar Ali, Jaan Praks, and Leonardo M. Reyneri. "Optimized Design and Thermal Analysis of Printed Magnetorquer for Attitude Control of Reconfigurable Nanosatellites." IEEE Transactions on Aerospace and Electronic Systems 56, no. 1 (February 2020): 736–47. http://dx.doi.org/10.1109/taes.2019.2933959.

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13

Feng, Guanhua, Chen Zhang, Heng Zhang, and Wenhao Li. "Theoretical and Experimental Investigation of Geomagnetic Energy Effect for LEO Debris Deorbiting." Aerospace 9, no. 9 (September 14, 2022): 511. http://dx.doi.org/10.3390/aerospace9090511.

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Space debris is increasingly problematic and needs active removal, especially in low Earth orbits (LEO). Paying for the vast cost of the disposal of debris from the situation is still inevitable even though pivotal technical hurdles have been overcome with the growing maturity of capturing and deorbiting methods. To this end, a novel geomagnetic energy (GME) propellant approach is firstly proposed to propel a spinning tethered spacecraft for LEO debris deorbiting, without the use of expendable fuel and a large-length tether. In this method, the time-cumulative effect of the interacted torque of the spacecraft’s electromagnet and geomagnetic field is used to accelerate the rotating system for GME storage, and the space momentum exchange from the angular momentum of system to the linear momentum of debris is introduced to deorbit the debris for GME release. Next, an on-orbit directional GME storage mechanism is built, and the corresponding two optimal strategies are put forward. Both theoretical and simulation results demonstrate that GME can be stored in the expected direction on any inclined LEO below 1000 km. Deorbiting kg-level debris can be accomplished within several orbital periods with the existing magnetorquer technology. Finally, proof-of-principle experiments of the GME effect are performed and elementarily validate the LEO GME utilization in space.

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14

Sun, Liang, Zhiwen Wang, Guowei Zhao, and Hai Huang. "Magnetic attitude tracking control of gravity gradient microsatellite in orbital transfer." Aeronautical Journal 123, no. 1269 (September 4, 2019): 1881–94. http://dx.doi.org/10.1017/aer.2019.112.

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ABSTRACTThe problem of the magnetic attitude tracking control is studied for a gravity gradient microsatellite in orbital transfer. The contributions of the work are mainly shown in two aspects: (1) the design of an expected attitude trajectory; (2) a method of the magnetic attitude tracking control. In orbital transfer, the gravity gradient microsatellite under a constant thrust shows complicated dynamic behaviours. In order to damp out the pendular motion, the gravity gradient microsatellite is subject to the the attitude tracking problem. An expected attitude trajectory is designed based on dynamic characteristics revealed in the paper, which not only ensures the flight safety of the system, but also reduces the energy consumption of the controller. Besides, the control torque produced by a magnetorquer is constrained to lie in a two-dimensional plane orthogonal to the magnetic field, so an auxiliary compensator is proposed to improve the control performance, which is different from existing magnetic control methods. In addition, a sliding mode control based on the compensator is presented, and the Lyapunov stability analysis is performed to show the global convergence of the tracking error. Finally, a numerical case of the gravity gradient microsatellite is studied to demonstrate the effectiveness of the proposed tracking control.

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15

Feng, Guanhua, Wenhao Li, and Heng Zhang. "Geomagnetic Energy Approach to Space Debris Deorbiting in a Low Earth Orbit." International Journal of Aerospace Engineering 2019 (February 7, 2019): 1–18. http://dx.doi.org/10.1155/2019/5876861.

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The space debris removal problem needs to be solved urgently. Over 70% of debris is distributed between the 500 km and 1000 km low Earth orbits (LEO), and existing methods may be theoretically feasible but are not the high-efficiency and low-consumption methods for LEO debris removal. Based on the torque effect of a static magnet interacting with the geomagnetic field, a new spin angular momentum exchange (SAME) method by geomagnetic excitation (without working medium consumption) for LEO active debris deorbiting is proposed. The LEO delivery capability of this method is researched. Two kinds of spin angular momentum accumulation (SAMA) strategies are proposed. Then through numerical simulation under the dipole model and International Geomagnetic Reference Field (IGRF11) model, the results confirm the physical feasibility and basic performance of the proposed method. The method can be applied to the regions of the LEO below 1000 km with different altitudes/inclinations and eccentricities, and with existent magnetorquer technology, only several days of preparation is required for about 104 m·kg mechanism-scale-debris-mass deorbiting, which can be used for deorbiting missions in debris-intensive areas (altitude≤1000 km); without consideration of external effects on the geomagnetic field distribution, it has the same deorbiting capability with that of the LEO below 1000 km when the altitude is over 1000 km. Besides, the method is characterized by explicit mechanism, flexible control strategy and application, and low dependence on the scale. Finally, the key technology requirements and future application of LEO active debris removal and on-orbit delivery by using SAME are prospected.

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16

Zheliabov, Petro, and Erik Lapkhanov. "Development of the methodological approaches for the attitude control system of the Earth remote sensing satellite in the conditions of the onboard equipment partial failures." EUREKA: Physics and Engineering, no. 5 (September 30, 2022): 77–90. http://dx.doi.org/10.21303/2461-4262.2022.002020.

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The spacecraft controllability of the angular motion is possible only with operability of the attitude and orbit control system (AOCS) of the spacecraft, sensors, actuators and the spacecraft power system. However, there is a rather significant probability of failure of this equipment during the operation of the spacecraft. This is especially observed after half of the spacecraft's lifetime or because of emergency situations. There is a problem which is connected with providing the maximum performance of the AOCS in case of partial failures of their actuators (reaction wheels (RW), magnetorquer rods (MGTR), etc.). Thus, the purpose of this work is the development and synthesis of special algorithms for spacecraft angular motion control in the emergency situations which are connected with RWs partial failures and restrictions of onboard electricity consumption. The approach of synthesis of this control algorithms is based on using mobile control methods which allow to reserve RWs by MGTRs. There are different variants of control loops depending on MGTRs turning on combinations. There were proposed two types of control switching functions: time-periodic and switching by deviation. Also was proposed a methodology of controller synthesis using these switching functions. Using this methodology and computer simulation, it was shown the possibility of providing angular nadir orientation and stabilization of the spacecraft with maximum 1−1.5 deg error in case of time-periodic switching functions implementation. Switching by deviation allows to reduce onboard electricity consumption for 25−30 % comparing with using time-periodic switching. However, the accuracy of stabilization significantly lower in case of switching by deviation. Considering these estimates, the corresponding methodological recommendations were formulated for use switching functions depending on emergency

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17

Ivanov, Danil, Mikhail Ovchinnikov, and Dmitry Roldugin. "Three-Axis Attitude Determination Using Magnetorquers." Journal of Guidance, Control, and Dynamics 41, no. 11 (November 2018): 2455–62. http://dx.doi.org/10.2514/1.g003698.

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18

Misra, Rahul, Rafał Wisniewski, and Alexander Zuyev. "Attitude Stabilization of a Satellite Having Only Electromagnetic Actuation Using Oscillating Controls." Aerospace 9, no. 8 (August 13, 2022): 444. http://dx.doi.org/10.3390/aerospace9080444.

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We consider the problem of attitude stabilization for a low Earth orbit satellite having only electromagnetic actuation. Such a satellite is not fully actuated, as the control torque is the cross-product of magnetic moment due to magnetorquers and the geomagnetic field. The aim of this work is to study whether oscillating controls can be designed such that a satellite actuated via magnetorquers alone can achieve full three-axis control irrespective of the position of the satellite. To this end, we propose considering oscillating feedback controls which generate the motion of the closed-loop system in the direction of appropriate Lie brackets. Simulation studies show that the proposed control scheme is able to stabilize the considered system.

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19

Curatolo, Andrea, Anton Bahu, and Dario Modenini. "Automatic Balancing for Satellite Simulators with Mixed Mechanical and Magnetic Actuation." Aerospace 9, no. 4 (April 16, 2022): 223. http://dx.doi.org/10.3390/aerospace9040223.

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Dynamic spacecraft simulators are becoming a widespread tool to enable effective on-ground verification of the attitude determination and control subsystem (ADCS). In such facilities, the on-orbit rotational dynamics shall be simulated, thereby requiring minimization of the external torques acting on the satellite mock-up. Gravity torque is often the largest among the disturbances, and an automatic procedure for balancing is usually foreseen in such facilities as it is significantly faster and more accurate than manual methods. In this note, we present an automatic balancing technique which combines mechanical and magnetic actuation by the joint use of sliding masses and magnetorquers. A feedback control is employed for in-plane balancing in which the proportional and integral actions are provided by moving the masses, while the derivative action is provided by the magnetorquers. Compared to an earlier implementation by the authors relying on shifting masses only, the novel approach is shown to reduce the in-plane unbalance by an additional 45% on average.

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20

Trégouët, Jean-François, Denis Arzelier, Dimitri Peaucelle, and Luca Zaccarian. "Static input allocation for reaction wheels desaturation using magnetorquers." IFAC Proceedings Volumes 46, no. 19 (2013): 559–64. http://dx.doi.org/10.3182/20130902-5-de-2040.00066.

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21

Ivanov, Danil, Rui Gondar, Uliana Monakhova, Anna Guerman, and Mikhail Ovchinnikov. "Electromagnetic uncoordinated control of a ChipSats swarm using magnetorquers." Acta Astronautica 192 (March 2022): 15–29. http://dx.doi.org/10.1016/j.actaastro.2021.12.014.

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22

Wisniewski, Rafal, and Jakob Stoustrup. "Periodic H2 Synthesis for Spacecraft Attitude Control with Magnetorquers." Journal of Guidance, Control, and Dynamics 27, no. 5 (September 2004): 874–81. http://dx.doi.org/10.2514/1.10457.

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23

Giulietti, Fabrizio, Alessandro A. Quarta, and Paolo Tortora. "Optimal Control Laws for Momentum-Wheel Desaturation Using Magnetorquers." Journal of Guidance, Control, and Dynamics 29, no. 6 (November 2006): 1464–68. http://dx.doi.org/10.2514/1.23396.

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24

Wang, Ping, and Yuri Shtessel. "Satellite Attitude Control Via Magnetorquers Using Switching Control Laws." IFAC Proceedings Volumes 32, no. 2 (July 1999): 8021–26. http://dx.doi.org/10.1016/s1474-6670(17)57368-1.

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25

Karpenko, S. O., M. Yu Ovchinnikov, D. S. Roldugin, and S. S. Tkachev. "One-axis attitude of arbitrary satellite using magnetorquers only." Cosmic Research 51, no. 6 (November 2013): 478–84. http://dx.doi.org/10.1134/s0010952513060087.

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26

Tregouet, Jean-Francois, Denis Arzelier, Dimitri Peaucelle, Christelle Pittet, and Luca Zaccarian. "Reaction Wheels Desaturation Using Magnetorquers and Static Input Allocation." IEEE Transactions on Control Systems Technology 23, no. 2 (March 2015): 525–39. http://dx.doi.org/10.1109/tcst.2014.2326037.

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27

Bolandi, H., and B. G. Vaghei. "Stable Supervisory-Adaptive Controller for Spinning Satellite using Only Magnetorquers." IEEE Transactions on Aerospace and Electronic Systems 45, no. 1 (January 2009): 192–208. http://dx.doi.org/10.1109/taes.2009.4805273.

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28

Monkell, Matthew, Carlos Montalvo, and Edmund Spencer. "Using only two magnetorquers to de-tumble a 2U CubeSAT." Advances in Space Research 62, no. 11 (December 2018): 3086–94. http://dx.doi.org/10.1016/j.asr.2018.08.041.

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29

Grøtte, Mariusz Eivind, Jan Tommy Gravdahl, Tor Arne Johansen, Jesper Abildgaard Larsen, Edgard Martinez Vidal, and Egidijus Surma. "Spacecraft Attitude and Angular Rate Tracking using Reaction Wheels and Magnetorquers." IFAC-PapersOnLine 53, no. 2 (2020): 14819–26. http://dx.doi.org/10.1016/j.ifacol.2020.12.1924.

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30

Celani, Fabio. "Quaternion versus Rotation Matrix Feedback for Spacecraft Attitude Stabilization Using Magnetorquers." Aerospace 9, no. 1 (January 4, 2022): 24. http://dx.doi.org/10.3390/aerospace9010024.

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The purpose of this paper is to compare performances between stabilization algorithms of quaternion plus attitude rate feedback and rotation matrix plus attitude rate feedback for an Earth-pointing spacecraft with magnetorquers as the only torque actuators. From a mathematical point of view, an important difference between the two stabilizing laws is that only quaternion feedback can exhibit an undesired behavior known as the unwinding phenomenon. A numerical case study is considered, and two Monte Carlo campaigns are carried out: one in nominal conditions and one in perturbed conditions. It turns out that quaternion feedback compares more favorably in terms of the speed of convergence in both campaigns, and it requires less energy in perturbed conditions.

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31

Celani, Fabio. "Spacecraft Attitude Stabilization Using Magnetorquers with Separation Between Measurement and Actuation." Journal of Guidance, Control, and Dynamics 39, no. 9 (September 2016): 2184–91. http://dx.doi.org/10.2514/1.g001804.

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32

Celani, Fabio. "Robust three-axis attitude stabilization for inertial pointing spacecraft using magnetorquers." Acta Astronautica 107 (February 2015): 87–96. http://dx.doi.org/10.1016/j.actaastro.2014.11.027.

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33

Ivanov, Danil Sergeevich, and Dmitry Sergeevich Roldugin. "Satellite magnetic attitude: Lyapunov control and determination using electromotive force in magnetorquers." Keldysh Institute Preprints, no. 30 (2018): 1–24. http://dx.doi.org/10.20948/prepr-2018-30.

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34

Miranda, Francisco. "Guidance Stabilization of Satellites Using the Geomagnetic Field." International Journal of Aerospace Engineering 2012 (2012): 1–9. http://dx.doi.org/10.1155/2012/231935.

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In the last years the small satellites have played an important role in the technological development. The attractive short period of design and low cost of them and the capacity to solve problems that are usually considered as problems to big and expensive spacecrafts lead us to study the control problem of these satellites. Active three-axis magnetic attitude stabilization of a low Earth orbit satellite is considered in this work. The control is created by interaction between the magnetic moment generated by magnetorquers mounted on the satellite body and the geomagnetic field. This problem is quite complex and difficult to solve. To overcome this difficulty guidance control is considered, where we use ε-strategies introduced by Pontryagin in the frame of differential games theory. Qualitative analysis and results of numerical simulation are presented.

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35

Narkiewicz, Janusz, Mateusz Sochacki, and Bartłomiej Zakrzewski. "Generic Model of a Satellite Attitude Control System." International Journal of Aerospace Engineering 2020 (July 23, 2020): 1–17. http://dx.doi.org/10.1155/2020/5352019.

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A generic model of a nanosatellite attitude control and stabilization system was developed on the basis of magnetorquers and reaction wheels, which are controlled by PID controllers with selectable gains. This approach allows using the same architectures of control algorithms (and software) for several satellites and adjusting them to a particular mission by parameter variation. The approach is illustrated by controlling a satellite attitude in three modes of operation: detumbling after separation from the launcher, nominal operation when the satellite attitude is subjected to small or moderate disturbances, and momentum unloading after any reaction wheel saturation. The generic control algorithms adjusted to each mode of operation were implemented in a complete attitude control system. The control system model was embedded into a comprehensive simulation model of satellite flight. The simulation results proved the efficiency of the proposed approach.

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36

Bolandi, H., and B. G. Vaghei. "Errata: Stable Supervisory-Adaptive Controller for Spinning Satellite using Only Magnetorquers [Jan 09 192-208]." IEEE Transactions on Aerospace and Electronic Systems 45, no. 2 (April 2009): 802. http://dx.doi.org/10.1109/taes.2009.5089564.

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37

Sadigh, Sevil M., Abdorreza Kashaninia, and Seyyed Mohammad Mehdi Dehghan. "Adaptive Fault Tolerant Attitude Control of a Nano-Satellite with Three Magnetorquers and One Reaction Wheel." Journal of Aerospace Engineering 35, no. 1 (January 2022): 04021113. http://dx.doi.org/10.1061/(asce)as.1943-5525.0001321.

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38

Wiśniewski, Rafal, and Jakob Stoustrup. "Periodic H 2 Synthesis for Spacecraft Attitude Determination and Control with a Vector Magnetometer and Magnetorquers." IFAC Proceedings Volumes 34, no. 12 (August 2001): 119–24. http://dx.doi.org/10.1016/s1474-6670(17)34072-7.

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39

Ovchinnikov, M. Yu, V. I. Penkov, D. S. Roldugin, and S. S. Tkachev. "Single axis stabilization of a fast rotating satellite in the orbital frame using magnetorquers and a rotor." Acta Astronautica 173 (August 2020): 195–201. http://dx.doi.org/10.1016/j.actaastro.2020.04.057.

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40

Yadegari, Hamed, Jalil Beyramzad, and Esmaeel Khanmirza. "Magnetorquers-based satellite attitude control using interval type-II fuzzy terminal sliding mode control with time delay estimation." Advances in Space Research 69, no. 8 (April 2022): 3204–25. http://dx.doi.org/10.1016/j.asr.2022.01.018.

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41

GONZALEZ JORGE, HIGINIO, NOELIA FARIÑAS ALVAREZ, and FERMIN NAVARRO MEDINA. "METROLOGICAL EVALUATION OF HELMHOLTZ FACILITY SETUP FOR TESTING OF MAGNETIC ATTITUDE DETERMINATION AND CONTROL SYSTEMS (ADCS) OF SMALL SATELLITES." DYNA 97, no. 3 (May 1, 2022): 267–73. http://dx.doi.org/10.6036/10380.

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The number of CubeSats launched to space has largely increased during last years. Any kind of satellite, even CubeSats, need to have some control on their attitude for pointing, i.e. antenna to Earth or solar panels to Sun. Among the available technology to control the attitude, CubeSats usually have cylindrical coils called magnetorquers, whose magnetic field interacts with the Earth’s magnetic field. CubeSat To verify its performance, it can be used a testing facility capable of simulating magnetic field along the satellite orbit. This work presents a metrological uncertainty analysis of a testing facility based on Helmholtz cage. Magnetic field uncertainty is obtained by performing uniformity, stability and gaussmeter calibration error tests. To simulate low Earth orbit satellite magnetic field conditions, the International Geomagnetic Reference Field model is used. The operation algorithms and software, developed using Python language, are validated by contrasting its results with Systems Tool Kit software, a commercial package, showing a maximum relative error of 0.03 %. Values from the testing facility are compared with in-situ measurements from Lume-1 satellite on-board gaussmeter, showing a maximum discordance lower than 50 mG. Keywords: spacecraft tests; attitude control system; Helmholtz facility; uncertainty evaluation; metrology

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Li, Junquan, Mark Post, Thomas Wright, and Regina Lee. "Design of Attitude Control Systems for CubeSat-Class Nanosatellite." Journal of Control Science and Engineering 2013 (2013): 1–15. http://dx.doi.org/10.1155/2013/657182.

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We present a satellite attitude control system design using low-cost hardware and software for a 1U CubeSat. The attitude control system architecture is a crucial subsystem for any satellite mission since precise pointing is often required to meet mission objectives. The accuracy and precision requirements are even more challenging for small satellites where limited volume, mass, and power are available for the attitude control system hardware. In this proposed embedded attitude control system design for a 1U CubeSat, pointing is obtained through a two-stage approach involving coarse and fine control modes. Fine control is achieved through the use of three reaction wheels or three magnetorquers and one reaction wheel along the pitch axis. Significant design work has been conducted to realize the proposed architecture. In this paper, we present an overview of the embedded attitude control system design; the verification results from numerical simulation studies to demonstrate the performance of a CubeSat-class nanosatellite; and a series of air-bearing verification tests on nanosatellite attitude control system hardware that compares the performance of the proposed nonlinear controller with a proportional-integral-derivative controller.

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Sedelnikov, Andry, Sergey Safronov, and Ekaterina Khnyryova. "The Rotational Motion Simulation of AIST Small Spacecraft Prototype Based on Current Values From Solar Battery Panels and on a Hardware-software Stand." International Journal of Mathematics and Computers in Simulation 15 (December 29, 2021): 165–68. http://dx.doi.org/10.46300/9102.2021.15.30.

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The functioning of space technique is associated with remote maintenance of operable state of onboard systems and software. If an element fails, problems arise in analyzing the state, ascertaining the reasons for the failure and restoring the element functional using available hardware-in-the-loop and algorithmic tools. The paper concentrates on currents analysis from solar battery panels of AIST small spacecraft in order to evaluate the parameters of the satellite's rotational motion after a significant break-down of the electrical battery. At the same time, the scientific equipment and onboard measurement instruments proved to be practically inoperative due to the lack of power supply. After the break-down of the electrical battery, magnetorquers and measurement instruments could not perform their function. A backup orientation system was not provided. The raw data for estimating the angular velocity vector was the current values from solar battery panels. However, in order to obtain an acceptable estimate of the angular velocity vector, more accurate current measurements are required than it implemented onboard the small spacecraft. To simulate the small spacecraft rotational motion and compare results with estimates obtained from telemetric data analysis, HIL US-03 hardware-software stand for simulation of the small spacecraft systems was used.

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Sedelnikov, Andry, Ekaterina Khnyryova, and Tatiana Ivashova. "Checking the correct operation of main measuring instruments on the flight model and prototype of AIST small spacecraft." MATEC Web of Conferences 234 (2018): 01007. http://dx.doi.org/10.1051/matecconf/201823401007.

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Flight model and prototype of AIST small spacecraft were launched in 2013 and 2014 respectively. They were used for carrying out various scientific researches, as well as testing the operation of AIST series small spacecrafts. They were operated in an uncontrolled flight and did not have a fullyfeatured orbital motion control system. However, it has been possible to reduce the angular velocity by magnetorquers. To obtain the control laws, the angular velocity of small spacecraft was estimated by means of magnetometer sensors. On the flight model of AIST small spacecraft, the angular velocity reduction mode was used twice. In the first case, the angular velocity was reduced, in the second case, it did not change significantly. On the prototype of AIST small spacecraft, the angular velocity reduction mode was used three times. In all three cases, a significant increase in the angular velocity of small spacecraft rotation was observed. The paper describes a test for checking the magnetometer sensors correct operation. Based on this test, conclusions about the measuring equipment correct operation are made. Ineffective control of the angular velocity of small spacecraft rotation is most likely due to failure to take account of impact of scientific and supporting equipment on the measuring instruments.

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Romero, Alessandro Gerlinger, and Luiz Carlos Gadelha De Souza. "Stability Evaluation of the SDRE Technique based on Java in a CubeSat Attitude and Orbit Control Subsystem." WSEAS TRANSACTIONS ON SYSTEMS 20 (January 29, 2021): 1–8. http://dx.doi.org/10.37394/23202.2021.20.1.

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In 2013, the STRaND (University of Surrey and Surrey Satellite Technology Ltd) and the PhoneSat (NASA) programs attracted the attention of the aerospace community applying commercial off-the-shelf smartphones in CubeSats. Both programs deployed CubeSats using smartphones based on Google's Android, in which application development is mainly based on Java programming language. Some of these CubeSats had actuators, e.g., STRaND-1 had three reaction wheels mounted in an orthogonal configuration to provide three-axis control, whereas PhoneSat 2.0 beta had magnetorquers to de-tumble the spacecraft. Taking into account a CubeSat that runs Android operating system (based on a smartphone), it is natural to evaluate the attitude and orbit control subsystem (AOCS) based on Java. Elsewhere, we shown State-Dependent Riccati Equation (SDRE) is a feasible non-linear control technique that can be applied in such CubeSats using Java. Moreover, we shown, through simulation using a Monte Carlo perturbation model, SDRE provides better performance than the PID controller, a linear control technique. In this paper, we tackle the next fundamental problem: stability. We evaluate stability from two perspectives: (1) parametric uncertainty of the inertia tensor and (2) a Monte Carlo perturbation model based on a uniform attitude probability distribution. Through the combination of these two perspectives, we grasp the stability properties of SDRE in a broader sense. In order to handle the uncertainty appropriately, we combine SDRE with H∞. The Nanosatellite Constellation for Environmental Data Collection (CONASAT), a CubeSat from the Brazilian National Institute for Space Research (INPE), provided the nominal parameters for the simulations. The initial results of the simulations shown that the SDRE controller is stable to ± 20% uncertainty in the inertia tensor for attitudes uniformly distributed and angular velocity up to 0.15 radians/second.

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Alpatov, Anatolii, and Erik Lapkhanov. "THE USE OF MOBILE CONTROL METHODS FOR STABILIZATION OF A SPACECRAFT WITH AEROMAGNETIC DEORBITING SYSTEM." System technologies 6, no. 125 (December 27, 2019): 41–54. http://dx.doi.org/10.34185/1562-9945-6-125-2019-04.

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The search for optimal control algorithms for spacecrafts is one of the key areas in rocket and space technology. Taking into account certain restrictions and requirements in a specific space mission, the selection of certain executive devices of the spacecraft is carried out and the corresponding control law is synthesized. One of such space missions is the providing of angular motion stabilization of a utilized spacecraft with aeromagnetic deorbiting system. The stabilization of spacecraft angular motion is needed for the orientation of aerodynamic element perpendicular to the vector of atmosphere dynamic flux with the aim of increasing of aerodynamic braking force. In this mission, the main optimization criterion is the minimization of the on-board electrical energy consumption which is needed for the control of angular motion. The original construction of the aeromagnetic deorbiting system consists of aerodynamic flat sails element and executive control devices with permanent magnets. However, not all spacecraft can be equipped with additional executive control devices with permanent magnets. That’s why with the aim of expansion of aeromagnetic deorbiting system application, using extra source of electromagnetic control executive devices is proposed in this research.The purpose of the article is the search of the control law which provides minimal consumption of electrical on-board energy by electromagnetic control executive devices during long-term deorbiting mission. For satisfying this criterion of optimization using of mobile control methods to orientate the spacecraft with aeromagnetic deorbiting system are proposed in this investigation. Computer modeling of orbital motion of spacecraft with aeromagnetic deorbiting system show the efficiency of using proposed mobile methods for angular motion control which realized by electromagnetic devices – magnetorquers. It has been showed that because of using mobile control method consumption of on-board electrical energy significantly less than with classical approach. The advantages and disadvantages have been determined.

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Tayebi, Javad, Chao Han, and Yuanjin Yu. "Disturbance observer backstepping sliding mode agile attitude control of satellite based on the anti-saturation hybrid actuator of the magnetorquer and magnetically suspended wheel." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, February 23, 2022, 095441002110696. http://dx.doi.org/10.1177/09544100211069699.

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This paper presents a new anti-saturation strategy to avoid external singularity, improve reliability, and fast maneuvers of a satellite by combining the novel three-dimensional magnetically suspended wheel (3-D MSW) actuator and magnetorquer. The MSW is the preferred actuator for agile maneuvering compared to the traditional control moment gyro due to frictionlessness, low vibration, and long lifetime. The anti-saturation strategy includes 3-D MSW arrangement with pyramid configuration and three orthogonal magnetorquers. A steering law is designed to distribute control torque between actuators by calculating command tilt angles of the 3-D MSW and magnetic dipoles of the magnetorquer. A nonlinear disturbance observer backstepping sliding mode controller is designed to control the rotor shaft of 3-D MSW for agile maneuvers and desaturate it with the magnetorquer despite high-frequency disturbances. Finally, simulation results of attitude control based on anti-saturation hybrid actuators demonstrate the effectiveness and accuracy of agile maneuvers. Compared with usual steering laws, simulation outcomes confirm the proposed method and desaturate the 3-D MSW system when it experiences a maximum workspace and does not have extra angular momentum.

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Ali, A., W. Chao, S. A. Khan, J. Tong, and L. M. Reyneri. "Embedded magnetorquer for the more demanding multi-cube small satellites." Aeronautical Journal, March 17, 2022, 1–13. http://dx.doi.org/10.1017/aer.2022.24.

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Abstract The recent development in the miniaturisation of small satellites and their subsystems has opened a new window of research for the universities around the globe. The low-cost, lightweight, small and flexible satellites have resulted in a broad range of multi-cube format small satellites, constructed from one-to-many adjoined cubes, having total mass between 1 and 10kg. The most challenging design part of the small satellites is to implant a large number of subsystems in a limited space. In order to resolve this issue, the designers are trying to shrink down the subsystem’s dimensions further. In this paper, a magnetorquer coil is designed and analysed for a 4U (4 units cube; 33 × 33 × 16.5)cm3 and 8U (8 units cube; 33 × 33 × 33)cm3 multi-cube small satellites, respectively. The coil is embedded in the six internal layers of an eight-layers printed circuit board (PCB). The designed magnetorquer system is fully reconfigurable and multiple coils configurations can be achieved by attaching them in series, parallel and hybrid arrangements. Due to embedded nature, the heat generated by the coil may damage the components mounted on the PCB outer surfaces. Therefore, thermal analysis is performed to ensure that the coil generated heat will not cross the PCB components temperature safety limits. All the possible combinations of the coils are analysed for current drawn, power consumption, heat dissipation, magnetic moment generation and resultant torque. A desired torque can be attained by using a particular coil configuration at the cost of specific amount of consumed power and PCB surface thermals.

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Jovanovic, N., B. Riwanto, P. Niemela, M. Rizwan Mughal, and J. Praks. "Design of magnetorquer based attitude control subsystem for FORESAIL-1 satellite." IEEE Journal on Miniaturization for Air and Space Systems, 2021, 1. http://dx.doi.org/10.1109/jmass.2021.3093695.

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Ali, Hassan, Qamar ul Islam, M. Rizwan Mughal, Rehan Mahmood, M. Rizwan Anjum, and Leonardo M. Reyneri. "Design and Analysis of a Rectangular PCB Printed Magnetorquer for Nanosatellites." IEEE Journal on Miniaturization for Air and Space Systems, 2020, 1. http://dx.doi.org/10.1109/jmass.2020.3029489.

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