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In the last decades, carbon-fiber-reinforced composite plates (CFRPs) have been widely employed for various structural applications on board spacecraft. Compared to metals, CFRPs present different mechanical properties related to the anisotropy of the fibers; this can strongly influence the response of CFRPs to hypervelocity impacts. In this work, four impact experiments performed on thin panels of CFRPs are presented to assess the influence of the impact velocity and of the impactor size on the fragmentation process; experimental data include fragment cumulative size and shape distributions. It is shown that size distributions strongly differ from aluminum plates in terms of the total number and shape. Significant differences are also shown on the shape distributions, which are mostly related to the presence of needle-shaped fragments.

This article presents an innovative mechatronic design of a high-accuracy pointing mechanism for orbital laser communication terminals.The system is based on a dual-stage architecture and is miniaturized to fit nanosatellite-class spacecraft, aiming to enable optical communication on small-size space platforms. The focus is on control design aspects and on the performance assessment of an experimental prototype under emulated external environmental disturbances.

Alba CubeSat UniPD is a student team of University of Padova with the aim to participate to the ESA Fly Your Satellite! (FYS!) programme and to launch for the first time at University of Padova a CubeSat made by students. The proposed mission has three independent objectives: (1) to collect in-situ measurements of the sub-mm space debris environment in LEO, (2) to study the micro-vibration environment on the satellite throughout different mission phases, (3) to do precise orbit determination through laser ranging and evaluate procedures for fast satellite Pointing, Acquisition and Tracking (PAT) from ground. The proposed technological experiments aim to obtain data that will enrich the current knowledge of the space environment and will provide precious information useful for the further development of some research projects currently performed at University of Padova. In order to reach the objectives, in these years the activities of the teams aimed to develop a 2U CubeSat equipped with three payloads. The first payload is an impact sensor that will be placed on one of the outer faces of the satellite and will be able to count the number of debris impacting the spacecraft thus being able to measure the energy/momentum transferred to the satellite. The second one is a Commercial Off The Shelf (COTS) sensor that measures the micro-vibrations experienced by payloads in a CubeSat in different mission phases. The third one consists in a number of COTS Corner Cube Retroreflectors that will be placed onboard the satellite. Thanks to this, Satellite Laser Ranging (SLR) will be done to collect data on the satellite range and range rate using a facility currently under development at University. This paper presents the mission objectives and motivations. In addition, the mission phases and the preliminary design of the CubeSat reached during the activities of the project are shown. Particular attention is given to the payloads which are the most challenging aspect of this project.

This paper presents the design and characterization of a miniature docking mechanism for nanosatellites. Potential applications are several, including servicing of orbital vehicles (e.g. refuelling, components replacement, deorbiting or reboosting) and assembly of large structures (e.g. telescopes, antennas). The mechanism responds to the constant demand of enabling technologies from the booming small satellites market. The developed system has a traditional probe–drogue configuration; it is equipped with a sensor to detect contact and a single servo-actuator to lock the connection. The simple, though effective, design fills a relevant gap in the field of nanosatellite technologies. Numerical simulations have been conducted to evaluate the dynamics of docking procedures and to estimate loads exchanged at contact. Experimental results validate the simulations and prove the high tolerance to angular and linear (lateral) misalignment.

This paper provides a general description of a new Collision Simulation Tool (CST) to model the consequences of orbital impacts involving large debris and satellites, with the aim to predict fragments distributions in case of sub-catastrophic and catastrophic impacts. The new tool is being developed in the framework of the ESA contract “Numerical simulations for spacecraft catastrophic disruption analysis”, carried out by the Center of Studies and Activities for Space CISAS “G. Colombo” of the University of Padova (prime contractor) and etamax space GmbH. The CST makes possible to model a large variety of collision scenarios involving complex systems, such as entire satellites with many design details included, and provides statistically accurate results with a computational effort orders of magnitude lower than hydrocodes. To this end, the simulation approach is based on a hybrid modelling strategy, in which every colliding object is described as a gross net of Macroscopic Elements (ME) representing spacecraft elementary building blocks. On one hand, individual fragmentation of Macroscopic Elements is addressed through the use of semi-empirical breakup models applied, at element level, only to those spacecraft parts which are involved in the collision. On the other hand, structural distortion, fracture, and separation of satellite broken parts are modelled with a discrete-element approach through the simulation of momentum transfer to Macroscopic Elements through the net, taking into account energy dissipation inside elements and across links. This paper provides a description of the CST simulation methodology and framework; preliminary results are also shown with respect to the tool validation process. Validation is done by comparing the tool predictions with empirical data from ground-based impact tests on simple targets (simple plates and Whipple Shields) as well as sub-scale spacecraft models. In the first case, the tool capability of predicting fragments distributions from plates and ballistic limit equations of shields is verified. In the second case, fragments distributions calculated by the software are compared with those derived from published experiments on a micro-satellite.

This paper presents the design and testing of LaserCube, a miniature optical communication terminal conceived for nano and microsatellites. The system architecture has been defined for both the downlink and intersatellite link version of the system. Then, a complete engineering model of LaserCube in its intersatellite link configuration has been developed and tested. It features (1) a dual stage pointing and tracking system based on a coarse pointing mechanism patented by Stellar Project, (2) an optical head with a full–duplex telecom channel with transmission and reception on the same wavelength for two–way links, (3) a transceiver section with telecom laser source and optical receiver and (4) the terminal control unit with onboard computer, actuator drivers and data interface. Experimental validation of the system is achieved through a laboratory simulation of an intersatellite link scenario with realistic dynamic disturbance coming from the host satellite attitude jitter.

Small satellite platforms are becoming very appealing both for scientific and commercial applications, thanks to their low cost, short development times and availability of standard components and subsystems. The main disadvantage with such vehicles is the limitation of available resources to perform mission tasks. To overcome this drawback, mission concepts are under study that foresee cooperation between autonomous small satellites to accomplish complex tasks; among these, on-orbit servicing and on-orbit assembly of large structures are of particular interest and the global scientific community is putting a significant effort in the miniaturization of critical technologies that are required for such innovative mission scenarios. In this work, the development and the laboratory testing of an accurate relative navigation package for nanosatellites compliant to the CubeSat standard is presented. The system features a small camera and two sets of LED fiducial markers, and is conceived as a standard package that allows small spacecraft to perform mutual tracking during rendezvous and docking maneuvers. The hardware is based on off-the-shelf components assembled in a compact configuration that is compatible with the CubeSat standard. The image processing and pose estimation software was custom developed. The experimental evaluation of the system allowed to determine both the static and dynamic performances. The system is capable to determine the close range relative position and attitude faster than 10 S/s, with errors always below 10 mm and 2 deg.

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.

TED (Tethered Electromagnetic Docking) is a system proposed by a group of researchers and students of the University of Padova for close rendezvous and docking between spacecraft. It consists in a small tethered probe ejected by the chaser, reaching the proximity of the target with a controlled deployment, and then magnetically guided by a receiving electromagnet mounted on it. Because of the generated magnetic field, alignment and mating are possible, then, as the tether is rewound, the chaser is able to dock with the target. To perform a preliminary verification of TED, three groups of students have been involved in the project and contributed to the evaluation of its critical technologies in reduced gravity: in the framework of ESA &ldquo, Drop your Thesis!&rdquo, 2014 and 2016 campaigns the experiments FELDs and STAR focused on the test of the tether deployment and control, while PACMAN, in the framework of ESA &ldquo, Fly Your Thesis! 2017&rdquo, parabolic flights campaign, tested proximity operations by means of electromagnetic interactions. In this paper, a description of TED concept and its development roadmap is presented, introducing the critical technologies tested by FELDs, STAR, and PACMAN experiments. The second part of the paper focuses on the educational outcomes of the three experiments, introducing statistics on (1) student participation, (2) scientific publication production, and (3) influence of the educational programs on the students&rsquo, career.

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.

This paper presents a novel docking mechanism to provide small spacecraft with the ability to join and separate in space. At today small satellites mating technologies have never been verified in space nor scaled to CubeSat size; the few proposed ports implement simple probe-drogue interfaces, with the limitation to dock only with different-gender ports, or androgynous geometries, with complex and non-axis-symmetric latches. The mechanism presented in this work overcomes the aforementioned drawbacks, using a hybrid port that can act both as probe and drogue. This paper presents the mechanism design and analysis, focusing on the port kinematics and on the load transmission during its actuation and docking procedures. Experimental verification allows us to validate numerical simulations and determine the operative range of the port in terms of alignment ranges allowing the solid joint creation. Transmitted loads always under 3 N are shown during the docking transient, while the port displays to manage misalignments up to 5 ° and 15 mm.

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.

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.

In the framework of a collaboration between the Center of Studies and Activities for Space of the University of Padova (CISAS-UPD) and the Japan Aerospace Exploration Agency (JAXA), 45 hypervelocity impact experiments were realised on Carbon Fibre Reinforced Plastic samples (CFRP) having nominal thickness of 2.3, 3.5 and 4.3 mm. The impact tests were made with aluminium spheres of 0.8, 1.5, 1.9, 2.3 and 2.9 mm, at velocities between 2 and 5 km/s. After a preliminary post-impact inspection of the targets, ballistic limit equations were developed using a new procedure which permits to estimate the uncertainty in the failure predictions starting from the measurement of the projectile entry crater. Then, in case of sample perforation and consequent ejection of fragments, high-speed shadowgraphs of the debris cloud were analysed to qualitatively describe the evolution of the impact damage, together with its basic features.

This paper reports a theoretical and numerical study aimed at increasing the operating efficiency of two-stage light-gas guns by appropriately changing their working conditions. In particular, a method is presented for increasing the projectile speed without any rise of the maximum breech pressure. The classic design theory of two-stage guns starts from the assumption that the highest velocity is reached in a gun which constantly maintains the maximum acceptable pressure at the base of the projectile during the full launching time. The main drawback of this working condition is that it may require an unfeasible rise of the gun maximum pressure, especially when very high muzzle speed is requested. To overcome this limitation, a new reference case different from the constant-base-pressure one is presented, based on a novel gas-dynamics solution that can be expressed in exact form if losses are not accounted for. According to such an approach, it is theoretically shown that the projectile base-pressure can be appropriately shaped (i) to improve the final speed without increasing the breech pressure or, in other terms, (ii) to achieve a given muzzle velocity with reduced maximum gas pressure. The analytical application of the new gas-dynamics condition showed the capability of obtaining a 1 km/s velocity improvement with no increase of the breech pressure or, alternatively, a pressure reduction up to 30% with no penalty on the model final speed. A numerical verification of the calculations was performed through the CISAS light-gas gun full numerical model, which includes real effects such as friction losses and heat transfer. Finally, an experimental verification of the numerical test case was attempted and a speed augmentation of 0.8 km/s with no increase in the breech pressure was confirmed in laboratory, highlighting the agreement with numerical predictions.

Impact processes influence the surface evolution of the solid bodies in the Solar System, as asteroids or comets. The better understanding of these phenomena can improve the interpretation of remote sensing data from forthcoming missions. In order to complement and to extend the available data on hypervelocity impacts on porous targets to ranges of velocity and physical conditions not yet explored, we focused our work on catastrophic fragmentation and cratering processes onto porous targets by means of high velocity impact experiments using a two-stage light-gas gun located at the impact facility of CISAS “G. Colombo” of the University of Padova. Tests have been performed on targets of different materials, e.g. glass ceramic foam, natural pumices, water ice, and different porosity (with density in the range from 0.35 to 1.07 g/cm 3 and porosity from 2% up to 65%). Results have been analysed and compared to published data. In particular, cratering morphology is studied as function of projectile velocity and density and then it is compared to empirical trend reported in Kadono [Kadono, T. Hypervelocity impact into low density material and cometary outburst. Planet. Space Sci., 47, 305–318, 1999.] with good agreement. Our tests results revealed that porosity seems to induce a rapid attenuation of the shock affecting energy propagation inside the target; craters generated by hypervelocity impacts on porous targets are smaller and much deeper than those of non porous targets. Fragments distribution resulting from disruption tests is analysed and the largest fragment mass has been studied as function of the collisional specific energy; results is compared to those of paper by Ryan et al. [Ryan, E.V., Davis, D.R., Giblin, I. A laboratory impact study of simulated edgeworth-Kuiper belt objects. Icarus 142, 56–62, 1999.]. Furthermore, numerical simulations have been performed by using Smooth Particle Hydrodynamics (SPH) and Lagrangian grid technique in order to test some available numerical models for a forthcoming more accurate numerical analysis that will be validated with the comparison with experimental results.

Soft robotics is a promising field and its application to space mechanisms could represent a breakthrough in space technologies by enabling new operative scenarios (e.g. soft manipulators, capture systems). Dielectric Elastomers Actuators have been under deep study for a number of years and have shown several advantages that could be of key importance for space applications. Among such advantages the most notable are high conversion efficiency, distributed actuation, self-sensing capability, multi-degree-of-freedom design, light weight and low cost. The big potentialities of double cone actuators have been proven in terms of good performances (i.e. stroke and force/torque), ease of manufacturing and durability. In this work the kinematic, dynamic and control design of a two-joint redundant robotic arm is presented. Two double cone actuators are assembled in series to form a two-link design. Each joint has two degrees of freedom (one rotational and one translational) for a total of four. The arm is designed to move in a 2-D environment (i.e. the horizontal plane) with 4 DoF, consequently having two degrees of redundancy. The redundancy is exploited in order to minimize the joint loads. The kinematic design with redundant Jacobian inversion is presented. The selected control algorithm is described along with the results of a number of dynamic simulations that have been executed for performance verification. Finally, an experimental setup is presented based on a flexible structure that counteracts gravity during testing in order to better emulate future zero-gravity applications.

This paper presents the results of hypervelocity impact experiments that were carried out at CISAS Impact Facility onto aluminum bumpers cooled down to −120 °C with liquid nitrogen and to −60 °C with solid carbon dioxide. The thickness of the targets was 0.8, 1, 2 and 3.17 mm, the diameter of the spherical projectiles was 1.5, 1.9, 2.3 and 2.9 mm and the impact velocity did span between 4 and 5 km/s. To establish if any temperature dependence exists in the bumpers’ impact response, two different features were analyzed: the hole size and the bumper protection capabilities. The latter property, that is related to the bumper capacity of producing debris cloud composed of fragments as fine and slow as possible, was assessed through observation of the damage patterns on witness plates and through measurements of the debris cloud tip velocity. Moreover, qualitative analyses of high-speed shadowgraphs representing the debris cloud evolution were performed. On one hand, it was found that low temperature has only minor influence on the hole diameter. On the other hand, the examination of shadowgraphs showed that the debris cloud structure varies with bumper temperature, even though it was not proved that such differences correspond to significant dissimilarities between damage patterns recorded onto witness plates.

Orbital robotics, due to the unfriendly environment (radiation, micro-gravity, thermal stresses, etc.) poses unique challenges to robot and robot algorithms, and sets the need for new and innovative autonomous systems. The design of servicing operations and devices is nowadays one of the most important research field in space robotics. Servicing operations range from regular inspection to the upgrade of components and re-fuelling. It is immediate to notice that, regardless of the operation to be carried out, the success is strictly linked to the way in which the chaser and the target satellites move and interact with respect to each other. The importance of relative motion for rendezvous and docking operations, calls for an approriate laboratory facility able to reproduce orbital conditions. This can be achieved only with a robotic structure that simulates the target and chaser's kinematics and dynamics. In this paper, a complete approach to the problem is presented, from the kinematic analysis to the modelling of the impact. In particular, a spring-dashpot model was chosen for the contact simulation, and a virtual-force control system has been adopted. Then, by considering the system's stability, we extracted the analytical expressions that link the performances of the facility with the range of orbital systems that can be simulated. Furthermore, with the aid of a SimMechanics® numerical model, we inspected the performances of three different control strategies for the movimentation of the robot.

This paper presents a novel set-up designed to investigate plastic deformation of a metal at high strain rates. The set-up is similar to conventional split Hopkinson pressure bar but the striker bar is eliminated and instead of it a spherical projectile, accelerated to high velocity with a two-stage light-gas gun, impacts and penetrates in a steel plate attached to the input bar. This results in propagation of plane waves in the bars and the sample. Impacts were carried out by aluminum spherical projectiles with diameter of 3.5 mm and velocity between 2 and 2.5 km/s onto front plates with thickness 10 mm and diameter of 55mm; the thickness of samples was between 3 and 7 mm. The experimental results were compared with 3D finite element simulation. Also, the effect of projectile velocity and sample thickness were investigated through experimental tests.

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to a total tally of approximately 7000 metric tons. Now, most of the cross-sectional area and mass (97% in LEO) is concentrated in about 4600 intact objects, i.e. abandoned spacecraft and rocket bodies, plus a further 1000 operational spacecraft. Simulations and parametric analyses have shown that the most efficient and effective way to prevent the outbreak of a long-term exponential growth of the catalogued debris population would be to remove enough cross-sectional area and mass from densely populated orbits. In practice, according to the most recent NASA results, the active yearly removal of approximately 0.1% of the abandoned intact objects would be sufficient to stabilize the catalogued debris in low Earth orbit, together with the worldwide adoption of mitigation measures. The candidate targets for removal would have typical masses between 500 and 1000 kg, in the case of spacecraft, and of more than 1000 kg, in the case of rocket upper stages. Current data suggest that optimal active debris removal missions should be carried out in a few critical altitude–inclination bands. This paper deals with the feasibility study of a mission in which the debris is removed by using a hybrid propulsion module as propulsion unit. Specifically, the engine is transferred from a servicing platform to the debris target by a robotic arm so to perform a controlled disposal. Hybrid rocket technology for de-orbiting applications is considered a valuable option due to high specific impulse, intrinsic safety, thrust throttle ability, low environmental impact and reduced operating costs. Typically, in hybrid rockets a gaseous or liquid oxidizer is injected into the combustion chamber along the axial direction to burn a solid fuel. However, the use of tangential injection on a solid grain Pancake Geometry allows for more compact design of the propulsion unit. Only explorative tests were performed in the past on this rocket configuration, which appears to be suitable as de-orbiting system of new satellites as well as for direct application on large debris in the framework of a mission for debris removal. The paper describes some critical aspects of the mission with particular concern to the target selection, the hybrid propulsion module, the operations as well as the systems needed to rendezvous and dock with the target, and the disposal strategy.

A new formulation of the spacecraft relative motion for a generic orbit is presented based on the orbital propagation method proposed by Peláez et al. in 2006 [1]. Two models have been developed. In the first model the method is applied to each spacecraft using a time synchronization of the system dynamical states. In the second model we employ a local orbital reference frame with a linearization of gravitational terms, apply the method to the formation center of mass and propagate the relative dynamics with respect to the center of mass reference orbit. The models are compared in terms of computational speed for the case of a bounded triangular formation

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 order to do risk assessments of debris impacts on unmanned spacecraft, it is necessary to investigate damage of honeycomb sandwich structures caused by debris impacts. However, the study of the honeycomb sandwich panel with CFRP face sheets has not been sufficiently performed. The purpose of this study is to investigate hypervelocity impact phenomena of CFRP-AL honeycomb sandwich structure. Hypervelocity impact tests were performed with a two-stage light gas gun at University of Padova. Three kinds of CFRP-AL honeycomb sandwich panels which are frequently used as a material of a spacecraft structure were tested. The cell size and the core thickness were varied. Aluminum spheres, 0.8 mm in diameter, were used as projectiles. The tests were performed at a velocity range between 2 and 5 km/sec. After the tests, the projectiles perforated all targets. The perforation holes on the panels were measured, and ultrasonic inspection was performed. The area of the perforation holes of the panel were increased with the impact velocity. The core size of the honeycomb core did not influence the relationship between the hole and the impact velocity. Impacts of the projectile on the foil of honeycomb cell caused heavy damage to a face sheet of the opposite side of the impact surface.

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.

The ReDSHIFT (Revolutionary Design of Spacecraft through Holistic Integration of Future Technologies) project was concluded on March 31, 2019. The 3-year project involved 13 European partners and was aimed at studying, implementing and testing novel solutions for space debris mitigation. The focus was on passive means to reduce the impact of Space Debris by prevention, mitigation and protection. The project was based on a synergy between theoretical and experimental aspects, such as: long term simulations, astrodynamics, passive de-orbiting devices, 3D printing, design for demise, hypervelocity impact testing, legal and normative issues. The paper presents an overview of the main results of ReDSHIFT, in an effort to highlight the holistic approach of the project covering different aspects of the space debris mitigation field.