Alessandro Francesconi

This work presents the development of the navigation and control subsystems of a guidance, navigation and control (GNC) system for controlling an autonomous satellite, called chaser, equipped with a redundant manipulator. In the study, the capture operations of a target spacecraft and its stabilization are considered. The control function employs a combined control strategy: the chaser satellite actuators and the robotic arm joint motors are degrees of freedom of the same control plant. Robust control methods are used to cope with uncertain, nonlinear dynamics of the chaser and of the complete chaser–target stack after capture. The developed navigation function is based on active or passive electro-optical sensors (i.e., passive cameras and/or LIDARs) for the relative pose determination. A numerical simulator capable of representing the dynamics of the system (chaser, manipulator, target) in orbit is developed in the MATLAB/Simulink environment to validate the developed GNC algorithms. The simulation tool allows to thoroughly assess the GNC system performance, accounting for all the relevant external disturbances and error sources. The simulator interfaces with a synthetic image generator (i.e., PANGU) for relative navigation performance assessment. This paper presents the testing of the GNC system in two relevant In-Orbit Servicing (IOS) scenarios: (1) the servicing of a large GEO platform and (2) the servicing of a small satellite in Low Earth Orbit (LEO).

Electromagnetic tethers of hundreds or thousands of meters have been proposed for maneuvring spacecraft in Low Earth Orbit, and in particular, for post-mission disposals. The debate on tether survivability to debris impact is still influencing further advances in the implementation of such technology because of the large area they expose to the debris environment; thin tape geometries have been proposed instead of round ones to increase the survivability to hypervelocity impacts. In this context, this paper introduces a new Ballistic Limit Equation (BLE) for thin tape tethers, derived from experimental results, numerical simulations, and literature data. The resulting equation is non-monotonic with respect to impact angle, presenting a minimum depending on the debris velocity and size; for high obliquities, the debris fragmentation triggered by shock waves propagating into the material reduces the damage. This feature allows to set a minimum particle diameter for risk assessment, excluding a significant part of the debris flux. The proposed BLE confirms the performance of thin tape tethers, with respect to round wires, due to their better ballistic response as well as their reduced cross-section at high-obliquity impacts.

Spacecraft fragmentation due to collisions with space debris is a major concern for space agencies and commercial entities, since in the next years the production of collisional fragments is expected to become the major source of space debris. Experimental studies have shown that the fragmentation process is highly complex and influenced by various factors, such as the satellite design, the material properties, the velocity and angle of the debris impact, and the point of collision (e.g., central, glancing, on spacecraft appendages). This paper summarizes the current state of research in spacecraft fragmentation, including the methods and techniques used to simulate debris impacts, the characterization of fragment properties and the analysis of the resulting debris cloud. It provides an overview of the main experiments performed, underlining the most critical issues observed. Moreover, it presents a set of experiments performed at the University of Padova and proposes some future directions for this research.

In the next years, the space debris population is expected to progressively grow due to in-space collisions and break-up events; in addition, anti-satellite tests can further affect the debris environment by generating large clouds of fragments. The simulation of these events allows identifying the main parameters affecting fragmentation and obtaining statistically accurate populations of generated debris, both above and below detection thresholds for ground-based observatories. Such information can be employed to improve current fragmentation models and to reproduce historical events to better understand their influence on the non-detectable space debris population. In addition, numerical simulation can also be used as input to identify the most critical objects to be removed to reduce the risk of irreversible orbit pollution. In this paper, the simulation of historical in-orbit fragmentation events is discussed and the generated debris populations are presented. The presented case-studies include the COSMOS-IRIDIUM collision, the COSMOS 1408 anti-satellite test, the 2022-151B CZ-6A in-orbit break-up, and a potential collision of ENVISAT with a spent rocket stage; for these events, results are presented in terms of cumulative fragments distributions and debris orbital distributions.

Simulation of in-space fragmentation events Lorenzo Olivieri, Cinzia Giacomuzzo, Stefano Lopresti, Alessandro Francesconi [button link="https://www.mrforum.