SITAEL (Italy)

Space exploration embodies humanity’s driver to transcend known boundaries and catalyses technological innovation, scientific advancements, and economic growth. As missions become increasingly complex, control theory emerges as a fundamental component, enhancing spacecraft navigation, operation, and adaptability in the dynamic space environment. This tutorial paper explores the pivotal role of control theory in advancing space exploration beyond the 2030s, emphasizing its contributions to autonomous decision-making, artificial intelligence, robust and resilient control, and the management of distributed systems. This paper outlines how advancements in control technologies will significantly enhance mission success and expand humanity’s presence in the solar system from both fundamental research and commercial viewpoints, with contributions from academia, space agencies, and industry. This paper presents a comprehensive overview on the traditional and intelligent technological solutions, addressing current limitations in space exploration while devising future possibilities.

Abstract The operation of a 20 kW magnetically shielded Hall effect thruster, with xenon and krypton as propellants, is numerically studied with a 2D (axial-radial) axisymmetric hybrid (particle-in-cell/fluid) code. Thrust efficiency with Kr is 2% to 6% lower than with Xe. Kr exhibits smaller propellant utilization, voltage utilization and divergence efficiencies. Contrary to what is observed in previous works, the current efficiency is higher with Kr. Apart from the dissimilar mass and electronic properties of Xe and Kr atoms, the widening of the ionization and acceleration zones with Kr, with respect to the Xe discharge, plays an important role in the Xe–Kr differences in performance. The level of plume velocity dispersion with both propellants is found similar due to two opposing effects: on the one hand, the stronger overlap of the ionization and acceleration zones with Kr leads to a larger dispersion on a per species basis; and, on the other hand, the smaller fraction of doubly-charged ions with Kr reduces the overall dispersion efficiency. Magnetic shielding is effective with both Xe and Kr as propellants, with slightly higher ion impact energy on the walls in the Kr discharge.

SITAEL, in collaboration with esc Aerospace, is developing an On-board Computer (OBC) in the framework of the ARTES Advanced Technology (AT) program. The unit is designed to be modular, by adhering to the SAVOIR generic OBC specifications and the cPCI Serial Space standard, and reconfigurable for future telecom (TLC) applications, being based on a large COTS reprogrammable SoC (System-on-a-Chip) FPGA.

High-power magnetically shielded Hall thrusters have emerged in recent years to meet the needs of the next-generation on-orbit servicing and exploration missions. Even though a few such thrusters are currently undergoing their late-stage development and qualification campaigns, many unanswered questions yet exist concerning the behavior and evolution of the plasma in these large-size thrusters that feature an unconventional magnetic field topology. Noting the complex, multi-dimensional nature of plasma processes in Hall thrusters, high-fidelity particle-in-cell (PIC) simulations are optimal tools to study the intricate plasma behavior. Nonetheless, the significant computational cost of traditional multi-dimensional PIC schemes renders simulating the high-power thrusters without any physics-altering speed-up factors unfeasible. The novel reduced-order ‘quasi-2D’ PIC scheme enables a significant reduction in the computational cost requirement of the PIC simulations. Thus, in this article, we demonstrate the applicability of the reduced-order PIC for a cost-efficient, self-consistent study of the physics in high-power Hall thrusters by performing simulations of a 20 kW-class magnetically shielded Hall thruster along the axial-azimuthal and radial-azimuthal coordinates. The axial-azimuthal quasi-2D simulations are performed for three operating conditions in a rather simplified representation of the thruster’s inherently 3D configuration. Nevertheless, we have resolved self-consistently an unprecedented 650 µs of the discharge evolution without any ad-hoc electron mobility model, capturing several breathing cycles and approximating the experimental performance parameters with an accuracy of 70%–80% across the operating conditions. The radial-azimuthal simulations, carried out at three cross-sections corresponding to different axial locations within the discharge channel, have casted further light on the evolution of the azimuthal instabilities and the resulting variations in the electrons’ cross-field mobility and the plasma-wall interactions. Particularly, we observed the development of a long-wavelength, relatively low-frequency wave mode near the exit plane of the thruster’s channel that induces a notable electron transport and a significant ion heating.

This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship.Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

Aims. Metis is the first solar coronagraph designed for a space mission and is capable of performing simultaneous imaging of the off-limb solar corona in both visible and UV light. The observations obtained with Metis aboard the Solar Orbiter ESA-NASA observatory will enable us to diagnose, with unprecedented temporal coverage and spatial resolution, the structures and dynamics of the full corona in a square field of view (FoV) of ±2.9° in width, with an inner circular FoV at 1.6°, thus spanning the solar atmosphere from 1.7R⊙to about 9R⊙, owing to the eccentricity of the spacecraft orbit. Due to the uniqueness of the Solar Orbiter mission profile, Metis will be able to observe the solar corona from a close (0.28 AU, at the closest perihelion) vantage point, achieving increasing out-of-ecliptic views with the increase of the orbit inclination over time. Moreover, observations near perihelion, during the phase of lower rotational velocity of the solar surface relative to the spacecraft, allow longer-term studies of the off-limb coronal features, thus finally disentangling their intrinsic evolution from effects due to solar rotation.Methods. Thanks to a novel occultation design and a combination of a UV interference coating of the mirrors and a spectral bandpass filter, Metis images the solar corona simultaneously in the visible light band, between 580 and 640 nm, and in the UV H ILyman-αline at 121.6 nm. The visible light channel also includes a broadband polarimeter able to observe the linearly polarised component of the K corona. The coronal images in both the UV H ILyman-αand polarised visible light are obtained at high spatial resolution with a spatial scale down to about 2000 km and 15 000 km at perihelion, in the cases of the visible and UV light, respectively. A temporal resolution down to 1 s can be achieved when observing coronal fluctuations in visible light.Results. The Metis measurements, obtained from different latitudes, will allow for complete characterisation of the main physical parameters and dynamics of the electron and neutral hydrogen/proton plasma components of the corona in the region where the solar wind undergoes the acceleration process and where the onset and initial propagation of coronal mass ejections (CMEs) take place. The near-Sun multi-wavelength coronal imaging performed with Metis, combined with the unique opportunities offered by the Solar Orbiter mission, can effectively address crucial issues of solar physics such as: the origin and heating/acceleration of the fast and slow solar wind streams; the origin, acceleration, and transport of the solar energetic particles; and the transient ejection of coronal mass and its evolution in the inner heliosphere, thus significantly improving our understanding of the region connecting the Sun to the heliosphere and of the processes generating and driving the solar wind and coronal mass ejections.Conclusions. This paper presents the scientific objectives and requirements, the overall optical design of the Metis instrument, the thermo-mechanical design, and the processing and power unit; reports on the results of the campaigns dedicated to integration, alignment, and tests, and to the characterisation of the instrument performance; describes the operation concept, data handling, and software tools; and, finally, the diagnostic techniques to be applied to the data, as well as a brief description of the expected scientific products. The performance of the instrument measured during calibrations ensures that the scientific objectives of Metis can be pursued with success.

There are a number of pressing problems mankind is facing today that could, at least in part, be resolved by space systems. These include capabilities for fast and far-reaching telecommunication, surveying of resources and climate, and sustaining global information networks, to name but a few. Not surprisingly, increasing efforts are now devoted to building a strong near-Earth satellite infrastructure, with plans to extend the sphere of active life to orbital space and, later, to the Moon and Mars if not further. The realization of these aspirations demands novel and more efficient means of propulsion. At present, it is not only the heavy launch systems that are fully reliant on thermodynamic principles for propulsion. Satellites and spacecraft still widely use gas-based thrusters or chemical engines as their primary means of propulsion. Nonetheless, similar to other transportation systems where the use of electrical platforms has expanded rapidly, space propulsion technologies are also experiencing a shift toward electric thrusters that do not feature the many limitations intrinsic to the thermodynamic systems. Most importantly, electric and plasma thrusters have a theoretical capacity to deliver virtually any impulse, the latter being ultimately limited by the speed of light. Rapid progress in the field driven by consolidated efforts from industry and academia has brought all-electric space systems closer to reality, yet there are still obstacles that need addressing before we can take full advantage of this promising family of propulsion technologies. In this paper, we briefly outline the most recent successes in the development of plasma-based space propulsion systems and present our view of future trends, opportunities, and challenges in this rapidly growing field.

Since 2003, the Microsatellite and Space Microsystems Lab at the University of Bologna (UniBo) has extended his research activities to the design of a ground segment for small satellites missions. In the framework of the European Student Earth Orbiter (ESEO), an ESA Education Office project for the development of a microsatellite mission, with SITAEL S.p.A. as the Industrial System Prime Contractor, the first-generation GS has been upgraded to support ESEO operations. UniBo was in charge for the design and development of the Mission Control Centre (MCC), the implementation of the primary ground station for telemetry and telecommand operations, and of the secondary one for the downlink of payloads data. ESEO was launched on December 3rd, 2018. Soon after launch, the activities planned for the Launch and Early Orbit Phase (LEOP) were forced to be held back due to the lack of success in commanding the spacecraft. After weeks of coordinated efforts among spacecraft operators, spacecraft engineers and ESA technical staff, LEOP activities could be resumed, thanks to the support of the radio amateur community and of ES5PC ground station in Tartu. Since then, new challenges are coming, which need to be faced for ESEO mission to advance further: this paper provides a thorough perspective of the achievements and lessons learnt during these months of operations.

This paper presents a "Medium Power" compact, light and low cost PCDU (Power Control and Distribution Unit) which has been designed as a modular and configurable equipment for Low Earth Orbit (LEO) missions with a maximum power capability of 1500 W. The PCDU supports a maximum of 4 independent power inputs (up to 75 V), and provides a 28 V unregulated power bus. Maximum Power Point Tracking (MPPT) is independently performed on each input; MPPT relies on buck regulators with direct drive capability, operating in `N+1' cold redundancy. The latter feature, together with the full cold redundancy of the Control Unit inside the PCDU, makes the device fully compliant with the Single Point of Failure free requirement. Moreover, the passive balancing managed by the PCDU guarantees the correct distribution of the cells charge, and contributes in improving the battery reliability. The unregulated bus is protected and distributed through nominal plus redundant LCL/RLCL and fuse protected lines for heaters and HDRMs. A cross-strapped architecture allows to control both nominal and redundant protected outputs from the nominal or redundant section of the unit, at the same time maximizing the PCDU availability. The PCDU, which is fully monitored, controlled and configured through the N+R CAN bus, can be easily tailored for various missions since its functionalities are segregated in different and specific stackable electronic boards (no backplane is used), and FPGAs are programmable. Finally, to balance the equipment radiation hardness and reliability figure versus costs, the PCDU has been designed with a mixed approach: COTS based supported by Sitael heritage [1] [2] and scheduled radiation tests, and the usage of MIL/Space qualified components for critical functions.