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The ever-increasing number of earth-orbiting spacecraft and related space junk is resulting in a dramatic rise in the risk of space debris impacting and damaging satellites and thereby negatively affecting the regular execution of several services provided by space-borne infrastructures. In the past years, the satellite market experienced a paradigm shift with the rise of small satellites and constellations formed by hundreds of satellites. It is anticipated that by the end of this decade, more than a thousand satellites per year will be launched, representing a potential market of more than $300 billion. With continued miniaturization of devices and the evolution of new mission requirements that rely on advanced sensor technology, future spacecraft will have an increasing density of devices and sensors. Moreover, a great research effort is required to improve the efficiency and reduce the weight of spacecraft shields. One route to achieve these goals is developing smart shields able to estimate the level of damage following the impact. In this context, the paper investigates a proof of concept based on the design, manufacture and testing of a measurement system, based on Thin Film Heat Flux Gauges (TFHFG), to assess the damage posed by orbital debris to the satellite shields upon Hyper Velocity Impacts (HVI). The system aims to measure the local increase in shield temperature, which is correlated to the kinetic energy of the debris. Following design and manufacturing, the proposed sensors were calibrated and mounted on a ductile aluminum alloy target, representative of the spacecraft shields, and subjected to a campaign of HVI tests. The results highlight that the signal is composed by the mechanical and thermal contribution, with a dominant mechanical factor.

Space missions that focus on On-Orbit-Servicing often involve the capture of a target satellite or space debris as a key event in the operational sequence. Space robots are among the most promising technologies to perform this task and mission architectures involving manipulators proved successful in orbit. Such mission scenarios are complex and imply a considerable risk of collision between the chaser and the target vehicles. For the sake of safety, it is of key importance to accurately position the capture interface w.r.t. the gripping point on the target, especially when relative distance is very low. During these phases, the guidance and control algorithms rely on pose measurements provided by navigation sensors that are generally able to reconstruct the target position and orientation only up to a certain distance between the camera and the target. To the purpose of having a complete control of the capture interface until contact, a new sensor based on a LED and a matrix of phototransistors has been developed. This paper describes the sensor and provides its preliminary characterization. A simple theoretical description of the sensor behavior is obtained from geometrical considerations, while experiments have been performed to measure the performances of the actual hardware implementation of the system.

A new statistical model for the characterisation of collision breakups in orbit is under development, as object of an ESA project titled "Exploiting numerical modelling for the characterisation of collision break-ups". The new model will be based on a large database of collision simulations, performed using a semi-empirical numerical tool, namely an improved version of the Collision Simulation Tool solver (CSTS), created by Centro di Ateneo di Studi e Attività Spaziali "Giuseppe Colombo" of the University of Padova (CISAS - Italy). The CSTS is used to reproduce the consequences of a collision between two orbiting structures, simulating their physical response at component level. The shape and model of the involved satellites/debris are taken into account, as well as the geometry of the impact. Small scale hypervelocity impact tests at component level are also performed to support the numerical tool validation and improvement, before the construction of the collision database. Exploiting the database of simulated collisions generated with the CSTS, the new statistical model aims to provide a fast and reliable tool for the prediction of the effects of a collision in orbit, taking into account the structure of the involved objects, their materials, as well as the collision geometry, thus overcoming some of the limitations of the currently available empirical and semi-empirical breakup models, like the de facto standard NASA breakup model (Johnson et al., 2001). As an example, these tools do not consider the objects' material and detailed design and do not take into account the effect of the impact point. On the basis of the input provided, the new model shall be able to work in basic mode, which only accounts for speed and masses, like the NASA model, or in advanced mode, taking into account also design details and collision geometry. A validation and comparison campaign is foreseen, also against the NASA model, on past in-orbit collisional events. In conclusion, the new tool shall be able to produce accurate information for many hypothetical collision scenarios in a relatively short CPU time. The final goal is to have a reliable computational tool that should be suitable to complement the standard NASA's breakup model for the predictions of the environmental consequences of a fragmentation in orbit. The content of the ongoing project activities and the partial results achieved at the time of the conference will be presented.

Recent plans for large constellations in Low-Earth Orbit have opened the debate on both their vulnerability and their influence on the already hazardous space debris environment. In fact, given that large constellations normally employ satellites of small size, there might be situations in which cm-size debris could have enough energy to cause fragmentation of a significant part of these spacecraft upon impact, while smaller debris could affect the functionalities of critical subsystems, even compromising the success of disposal operations planned at end-of-life. In this context, this paper investigates: (1) collisions with large objects that could initiate the fragmentation of a significant part of the satellite, and (2) impacts with small debris that might perforate the spacecraft hull thus causing relevant performance/functionality degradation. These two points are merged in a simple statistical tool for risk assessment, which analyses the effects of the main parameters of the constellations on its vulnerability (i.e. operational life, number of satellites, spacecraft cross section, satellites reliability). In more details, the tool relates impact probability (for both small and large debris) to the ballistic response of spacecraft structures and protections, defining the critical configurations that might compromise the expected disposal operations. This method requires a limited knowledge of the spacecraft internal layout, as it is based on a statistical analysis of impact damage instead of a complete evaluation of the vulnerability of each subsystem. In parallel, non-debris related failures are also investigated and statistic models of spacecraft reliability characteristic are proposed. Among the results, it is shown that reducing the lifetime of individual satellites in a constellation might improve the success rate of post-mission disposal, thanks to the reduction of the spacecraft exposure to the space environment with the consequential degradation of its performance. On the other hand, reducing the lifetime would seriously affect the debris environment: the increase in traffic in the most crowded altitudes would be not counterbalanced by the higher post mission disposal success rate, causing an overall increase of the total number of uncontrolled resident objects.

Honeycomb sandwich panels (HC/SP) are the most common used structures for space vehicle. Under the threat of meteoroid and space debris, the distribution of the fragments produced in a hypervelocity impact event on HC/SP is critical to the vulnerability assessment of space vehicle. CISAS developed an engineering model to describe fragments clouds propagating inside spacecraft in consequence of space debris impact on HC/SP. In this model, the effect of the honeycomb core was modeled by an empirical corrective factor, which was not related to the physical of the impact. To improve this model, a new model to describe the effect of the honeycomb core was developed. In the new model, the honeycomb core was equaled to multi-parallel thin plates, which can represent the discontinuity of honeycomb core without complex boundary. Based on the knowledge of hypervelocity impact on a simple thin plate and approximation supported by numerical simulation results, the model was deduced. The coefficient of the model was fitted by the numerical simulation results.

The main limitation of the currently adopted method for predicting spacecraft catastrophic fragmentation due to a collision with large debris is the total independence of the critical value of the energy-to-target mass ratio from both the satellite configuration and the impact point; in fact these two issues are not accounted for by the classical 40 J/g rule. To go beyond this limitation, the method proposed in this paper evaluates the distribution of impact energy into the system using the mechanical properties of the structural parts and the knowledge of the impact location. In this way, it becomes possible to predict how impact energy is partitioned among some selected macroscopic structural parts, each of them is finally evaluated versus its own minimum value of impact energy for which the part is fragmented (shattering threshold). Energy partition is performed by solving a system of equations written according to Statistical Energy Analysis (SEA). The paper describes in detail the proposed energy-partition method and presents its application to a geometrical representative model of a spacecraft subject to impact at different points. Results are finally compared to those obtained by the application of the classical 40 J/g rule. It is shown that the evaluation of spacecraft disintegration is highly influenced by the impact point and the structural properties of the components.

The H2020 ReDSHIFT project aims at finding passive means to mitigate the proliferation of space debris. This goal is pursued by a twofold research activity based on theoretical astrodynamics, computer simulations and the analysis of legal aspects of space debris, coupled with an experimental activity on advanced additive manufacturing (3D printing) applied to the production of a novel small satellite. Several different aspects related to the design and production of a debris compliant spacecraft are treated, including shielding, area augmentation devices for deorbiting (solar and drag sails) and design for demise. A strong testing activity, mainly based on design for demise wind tunnel experiments and hypervelocity impacts is performed as well. The main results obtained so far in the project are outlined.

Investigation of the effects of passive deorbiting devices (such as solar and drag sails) on the space debris environment. the key question to be answered is: are these devices a benefit or a risk for the growing space debris population?

ReDSHIFT will address barriers to compliance for spacecraft manufacturers and operators represented by requirements and technologies for de-orbiting and disposal of space objects. This will beachieved through a holistic approach that takes into account opposing and challenging constraints for the safety of the human population, when debris objects re-enter the atmosphere, designed fordemise, and for their survivability in the harsh space environment, while on orbit. ReDSHIFT will take advantage of disruptive opportunities offered by 3D printing to develop highly innovative, low-cost spacecraft solutions, exploiting synergies with electric propulsion, atmosphericand solar radiation pressure drag, and astrodynamical highways, to meet de-orbit and disposal needs, but which are also designed for demise. RedShift will design structures to enhance spacecraft protection, by fracture along intended breakup planes, and re-entry demise characteristics. These structures will be subjected to functional tests as well as specific hypervelocity impact tests and material demise wind tunnel tests to demonstrate the capabilities of the 3D printed structures. At the same time, novel and complex technical, economicand legal issues of adapting the technologies to different vehicles, and implementing them widely across low Earth orbit will be tackled through the development of a hierarchical, web-based toolaimed at a variety of space actors. This will provide a complete debris mitigation analysis of a mission, using existing debris evolution models and lessons learned from theoretical and experimentalwork. It will output safe, scalable and cost-effective satellite and mission designs in response to operational constraints. Through its activities, ReDSHIFT will recommend new space debris mitigationguidelines taking into account novel spacecraft designs, materials, manufacturing and mission solutions.

Electrodynamic tethers represent one of the possible means to de-orbit defunct satellites from Low-Earth-Orbit at end of life. However, tethers survivability to orbital debris impacts is still debated because of the large area they expose to the space environment. Recently, increasing consideration has been given to thin-tape tether geometries, whose response to space debris threat is believed to be better than that of round wires. This paper describes a new Ballistic Limit Equation (BLE) applicable for thin tapes, to go beyond previous investigations referring at most to the impact resistance of round-wires (furthermore neglecting the damage dependence from the impact velocity and angle). In this paper, a new approach for BLE derivation is presented, which combines experimental results (in total 24 impact tests) and numerical simulations (in total 112 runs). The resulting BLE is non-monotonic with respect to the impact angle and presents a minimum at certain values of the impact obliquity, depending from the debris size and speed. In other words, the minimum particle diameter which is just able to cut a tape at a given velocity decreases with increasing impact obliquity up to a certain angle above which the damage is reduced due to early debris fragmentation triggered by shock waves propagating into the material. Notably, it has been observed that there is a minimum value of debris velocity v* below which no critical damage is possible and, furthermore, there is a minimum velocity-dependent value d* of debris diameter below which no critical damage is possible. This feature of BLE is extremely important, since it sets a minimum particle diameter for risk assessment and thus excludes a large part of the flux from risk computations. In conclusion, the newly-developed BLE confirms that thin tapes are significantly more resistant than round wires of equivalent cross-section; this is due to the intrinsic ballistic response of tapes, not only to their reduced cross-section at high impact obliquity.

Electrodynamic tethers (EDT) represent one of the possible means to de-orbit defunct satellites from Low-Earth-Orbit at end-of-life. However, considering the large area exposed to the space environment and the consequent high number of debris impacts per unit time, a high tether survivability to orbital debris is of primary importance. This paper provides an estimation of the number of fatal impacts per unit time and per unit length on a tape tether, using for the first time an experimental ballistic limit equation that was derived for tapes and accounts for the effects of both the impact velocity and impact angle. It has recently been shown that, tape tethers, as opposite to round wires, are more resistant to space debris impacts. It is shown that considering a tape tether with cross section 45 mm × 50 μm , the number of critical events due to impact with non-trackable debris is always less than 0.01/yr/km, being maximum for orbit inclination of i = 90 ° .

All spacecraft in Earth orbit are exposed to the risk of impact with micrometeoroids and orbital debris. When such particles have enough energy to penetrate the hull of the vehicle, clouds of fragments are ejected into spacecraft and they can eventually compromise the functionality of various components encountered in their flight path. Knowledge of the clouds׳ properties (e.g. fragments mass and velocity) is therefore a key factor to obtain accurate predictions of the response of interior equipment to space debris threat. However, generation and evolution of debris clouds from hypervelocity impact is a complex phenomenon governed by a large number of parameters, and existing models mostly refer to fragments originated by impact on simple aluminum plates only, while the models available for sandwich panels do not provide information on the fragments mass. In such context, this paper presents an engineering model describing debris clouds created by space debris impacts on honeycomb sandwich panels representative of satellites structural bodies. The model describes the cloud as a cone with one large central fragment and several finely fragmented particles which behave like a jet of dust, and it provides a set of empirical equations predicting the cone geometry, the velocity and mass of the central fragment, as well as the total mass ejected downrange. The proposed equations are derived from analogous formulas for debris clouds originated by impacts on simple aluminum plates, by applying proper corrections to account for different materials effects and different behavior of sandwich panels compared to plates of same material. The model is finally evaluated by comparing its predictions with few experimental data on triple wall structures (sandwich panel plus internal equipment cover plate), where debris clouds exiting the panels’ rear skin are used to assess the failure of the third wall. In this latter case, the proposed model is also compared with the SRL equations.

This paper discusses the impact behavior of a self-healing ionomeric polymer and compares its protection capability against space debris impacts to that of simple aluminium-alloy bumpers. To this end, 14 impact experiments on both ionomer and Al-7075-T6 thin plates with similar surface density were made with 1.5 mm aluminium spheres at velocity between 1 and 4 km/s. First, the perforation extent in both materials was evaluated vis-a-vis the prediction of well known hole-size equations; then, attention was given to the damage potential of the cloud of fragments ejected from the rear side of the target by analysing the craters pattern and the momentum transferred to witness plates mounted on a ballistic pendulum behind the bumpers. Self-healing was completely successful in all but one ionomer samples and the primary damage on ionomeric polymers was found to be significantly lower than that on aluminium. On the other hand, aluminium plates exhibited slightly better debris fragmentation abilities, even though the protecting performance of ionomers seemed to improve at increasing impact speed.

This paper presents a case study of Micrometeoroids and Orbital Debris risk assessment for a spacecraft flying in an orbit close to that of the Italian Cosmo-Skymed constellation. The aim of the analysis was to calculate the failure flux impinging on the satellite external shell, taking into account both geometry and materials of satellite surfaces. Furthermore the analysis included the evaluation of the contribution to debris population at the selected orbit of the fragments produced by a Chinese Anti-SATellite experiment, which caused the catastrophic break-up of the satellite Fengyun 1C in January 2007. A first computation was carried out using ESABASE2/Debris v.1.4.2. This software made it possible to perform geometrical analysis of a satellite subjected to a given debris environment, but the most up to date available models, ORDEM2000 and MASTER2001, gave significantly different results. An independent procedure for risk assessment analysis was implemented to further analyse such issue and to provide damage equation adequate to represent the behaviour of the selected structural aluminium honeycomb sandwich panels covered by Multi-Layer Insulation. Debris fluxes were calculated applying MASTER2005 and ORDEM2000 environment models, then results were compared to those of ESABASE2. Failure fluxes were calculated implementing special damage equations for honeycomb structures available from the open technical literature. The expected flux contribution of catalogued debris from the Chinese Anti-SATellite (ASAT) experiment was estimated independently using the code SDIRAT (Space Debris Impact Risk Analysis Tool) developed at the Institute of Information Science and Technologies (ISTI).

In order to analyze the propagation of shocks due to hypervelocity impact of micrometeoroids and space debris on spacecrafts, it was necessary to analyze the high-frequency shock-propagation-dynamic on complex structures representative of real spacecraft structure. Objective of this research is the GOCE satellite due to its highly accurate accelerometer very sensitive to the micro-vibration environment. After a detailed analysis of the most-probable shock-propagation-path within the satellite, several representative targets have been designed. Then an extensive test campaign has been conducted on these targets exploring a wide range of impact conditions. As a result, a database was established which correlates the impact conditions in the experimental range (0.6–2.3 mm projectiles at 2.5–5 km/s) with the shock spectra on selected locations on various types of structural models. The behaviour of structural joints was also analyzed under shock conditions that are not normally reproduced by pyroshock testing. The database represent a fundamental tool in order to validate the numerical analysis that will be used to assess the vibration environment, due to in-orbit micrometeoroids and space debris at the accelerometer location. This paper presents the test procedure applied and the general results achieved.

In recent years, innovative processes such as additive manufacturing are finding application in the satellites industry, allowing solutions whose implementation could have been difficult or expensive using traditional manufacturing techniques. In the framework of the EU H2020 ReDSHIFT project, new 3D printed shields were developed for microsatellites protection: (1) a multi-shock panel comprising four equally-spaced 3D printed bumper layers, (2) a two-walls panel with outer bumper layers sandwiching a corrugated core, and (3) a hybrid panel with three equally-spaced bumpers, and a fourth corrugated bumper sandwiched between two of them. These shields were evaluated with respect to their resistance to hypervelocity impacts of mm-size debris and to radiation fluxes expectable in the LEO environment. This paper presents the results of hypervelocity impact and radiation tests carried out at the University of Padova on the proposed panels. On one hand, debris-shielding properties were evaluated with respect to the damage on targets and witness plates placed behind them. Debris clouds resulting from panels perforation were also observed qualitatively through high-speed imaging. On the other hand, radiation-shielding capabilities were assessed using a high-energy proton beam - protons are one of the most abundant ionizing particles in low orbits - and a combination of radiation sensors and relevant electronic chips. In particular, RADFETs were used to measure the amount of total ionizing dose with and without the interposition of the shields. In addition, Flash memories exposed to the proton beam, again with and without the shields, were used to directly evaluate radiation effects on chips relevant for space applications.