Attitude motion and sensor bias estimation onboard the SiriusSat-1 nanosatellite using magnetometer only

Ivanov, Danil; Roldugin, D. S.; Tkachev, S. S.; Ovchinnikov, M Y; Zharkikh, R. N.; Kudryavtsev, A. O.; Bychek, Mikhail

doi:10.1016/j.actaastro.2021.07.038

Cited by 7 publications

(2 citation statements)

References 21 publications

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“…This allowed reaching a high estimation accuracy, but it required a computational load and minimum sensor performances that are not suitable for nano-satellites. In these regards, the problem of attitude determination in nano-satellite was recently investigated by Fei [16], Li [17], and Ivanov [18]. These research works exploited miniaturized sensors; however, the obtained accuracy was in the range of 1 deg to 10 1 deg, and they did not provide MIL or HIL test results.…”

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confidence: 99%

An Effective Sensor Architecture for Full-Attitude Determination in the HERMES Nano-Satellites

Colagrossi

Lavagna

Bertacin

2023

Sensors

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The High Energy Rapid Modular Ensemble of Satellites (HERMES) is a constellation of 3U nano-satellites for high energy astrophysics. The HERMES nano-satellites’ components have been designed, verified, and tested to detect and localize energetic astrophysical transients, such as short gamma-ray bursts (GRBs), which are the electromagnetic counterparts of gravitational wave events, thanks to novel miniaturized detectors sensitive to X-rays and gamma-rays. The space segment is composed of a constellation of CubeSats in low-Earth orbit (LEO), ensuring an accurate transient localization in a field of view of several steradians exploiting the triangulation technique. To achieve this goal, guaranteeing a solid support to future multi-messenger astrophysics, HERMES shall determine its attitude and orbital states with stringent requirements. The scientific measurements bind the attitude knowledge within 1 deg (1σa) and the orbital position knowledge within 10 m (1σo). These performances shall be reached considering the mass, volume, power, and computation constraints of a 3U nano-satellite platform. Thus, an effective sensor architecture for full-attitude determination was developed for the HERMES nano-satellites. The paper describes the hardware typologies and specifications, the configuration on the spacecraft, and the software elements to process the sensors’ data to estimate the full-attitude and orbital states in such a complex nano-satellite mission. The aim of this study was to fully characterize the proposed sensor architecture, highlighting the available attitude and orbit determination performance and discussing the calibration and determination functions to be implemented on-board. The presented results derived from model-in-the-loop (MIL) and hardware-in-the-loop (HIL) verification and testing activities and can serve as useful resources and a benchmark for future nano-satellite missions.

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confidence: 99%

An Effective Sensor Architecture for Full-Attitude Determination in the HERMES Nano-Satellites

Colagrossi

Lavagna

Bertacin

2023

Sensors

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“…The work in [60] builds on [46] and considers the case where the magnetometer bias is varying as the residual magnetic dipole varies. The estimation error of the magnetometer bias was within 8%.…”

Section: Magnetometer Bias and Magnetic Attitude Determinationmentioning

confidence: 99%

Magnetic Disturbance Rejection in Spacecraft using Electromagnetic Propulsion

Jane

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<p><strong>Electric propulsion drives are the future of long-life and high velocity space missions due to their superior mass efficiency and reliability. In particular, applied field magnetoplasmadynamic thrusters (AF-MPDTs) are the class of electric propulsion that have the largest thrust and could give humanity access to greater depths of our solar system. However, control aspects of AF-MPDTs have not yet been investigated and potential problems may arise if this type of thruster is used in low Earth orbit. We have shown that the interaction of the geomagnetic field with the strong AF-MPDT magnetic field will cause unwanted magnetic disturbance torques on a spacecraft while it is manoeuvring. The spacecraft will gain angular momentum per orbit from these disturbance torques, for all orbital inclinations in low Earth orbit. For a 100 kg spacecraft, the reaction wheels required for momentum management are too large from a mass perspective to compensate the magnetic disturbance torques. Using the analysis of magnetic disturbance torques, a control solution is proposed with an Extended Kalman Filter to estimate the magnetic disturbance and a PD (proportional derivative) based magnetorquer control law to compensate for the estimated disturbance as well as maintain the attitude orientation required for an orbital manoeuvre. We have shown that a spacecraft using an electromagnetic thruster can be controlled in low Earth orbit for inclinations above 50 degrees, but control characteristics become poor for near equatorial orbits due to controllibility issues relating to the homogeneity of the geomagnetic field and limitations on magnetorquer properties.</strong></p>

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