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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 paper presents the experimental verification of the pointing and tracking system of the flight model of an optical communication terminal for small satellites. The terminal, called LaserCube, fits in two CubeSat units and is conceived to operate onboard spacecraft starting from the 6 U (CubeSat units) form factor. The terminal is provided with a dual stage pointing system, composed of a coarse pointing mechanism used to orient the optical head and a fast-steering mirror for the fine pointing of the received and transmitted lasers. The coarse pointing mechanism is based on the parallel platform configuration and alone can provide [Formula: see text] rms pointing accuracy when operating in conjunction with a beacon laser sent by a remote terminal and used for feedback. The control algorithm operates the two pointing stages synergically, targeting an overall pointing accuracy of [Formula: see text] rms while rejecting the disturbances coming from the satellite bus. The pointing system of the LaserCube flight model has been calibrated and experimentally verified in laboratory conditions that replicate real operational scenarios for what concerns the beacon signal attenuation and the disturbances generated by the host satellite. A dedicated setup was used to directly measure the pointing error and asses the system performance.

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.

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.

Small satellites are increasingly exploited for the realization of low Earth orbit (LEO) constellations for Earth imaging, global connectivity and data relay. Optical crosslinks between satellites can drastically increase the constellation capability to transmit data to ground stations with reduced latency and increased total throughput. One critical aspect of optical crosslink is the acquisition phase, which requires that each satellite scans its field of view looking for the beacon laser transmitted by the partner terminal before starting communication. In this paper, the problem of beacon acquisition for optical crosslinks is analyzed, with focus on the small satellite case. Numerical simulations and a simplified, yet representative laboratory experiment are carried out to evaluate and compare two different scanning techniques quantitatively.

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.

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.

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.

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.

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?

This paper provides an overview of the ARCADE-R2 experiment, a technology demonstrator that aimed to prove the feasibility of small-scale satellite and/or aircraft systems with automatic (a) attitude determination, (b) control and (c) docking capabilities. The experiment embodies a simplified scenario in which an unmanned vehicle mock-up performs rendezvous and docking opera- tions with a fixed complementary unit. The experiment is composed by a supporting structure, which holds a small vehicle with one translational and one rotational degree of freedom, and its fixed target. The dual system features three main custom subsystems: a relative infrared navigation sensor, an attitude control system based on a reaction wheel and a small-scale docking mechanism. The experiment bus is equipped with pressure and temperature sensors, and wind probes to monitor the external environmental conditions. The experiment flew on board the BEXUS 17 stratospheric balloon on October 10, 2013, where several

In this paper design, implementation, laboratory and microgravity tests of a space tether length and length rate measurer are discussed. The instrument presented was part of an experimental setup consisting in a new concept of space tether deployer with retrieval capability, which was designed and built in the framework of the Drop Your Thesis! 2016 educational programme, held by the European Space Agency Education Office. The instrument working principle is based on optical sensors and is inspired to similar instruments used on SEDS and YES2 space missions. The entire setup was first tested in laboratory on a low friction rail, then in microgravity conditions at the ZARM Drop Tower. The results of the test campaign proved the functionality of the measurement instrument which allowed the experiment to meet the prefixed goals.

Over the last years, nanosatellites based on the CubeSat standard have been increasingly exploited by both academic institutions and commercial companies. The low cost and short development time of CubeSats make such miniature spacecraft interesting for a variety of applications. A particularly appealing future field of application for nanosatellites is On-Orbit Servicing of existing orbital assets. A number of mission concepts foresee the employment of miniature spacecraft to perform a variety of operations on orbital vehicles, ranging from inspection and monitoring to assembly and repairing. In this framework, the development of technologies that enable proximity navigation and/or docking between a nanosatellite chaser and a target object is of interest. In particular, miniaturized navigation sensors for the estimation of relative position and attitude of the involved vehicles are required. The work presented here focuses on the development of an IR-based relative navigation sensor compatible with CubeSat standard nanosatellites. The system estimates the relative pose and position of the target by taking images of the object, which is equipped with retro-reflecting fiducial markers illuminated by an array of IR LEDs on the chaser. The system architecture and operation are described, and preliminary laboratory test results are presented.

A growing interest within the space community focuses on robotics due to the large number of possible applications in many mission scenarios. On-Orbit Servicing (OOS) is arguably the most appealing implementation of space automatic systems. In several cases, OOS requires the capture of orbital objects, which is a complex and risky operation that can be successfully performed by robotic manipulators. Soft robotics, in particular, seems to be suitable for such applications given its intrinsic compliance to the operative environment. Devices based on Dielectric Elastomers (DE) can be employed for the implementation of soft robotic systems and showed promising performances. The introduction of DEs to orbital systems would represent a breakthrough in space technologies. In addition, space conditions could further advantage DE robotics, given the reduced environmental loads experienced and the longer times for operations. Nevertheless, Dielectric Elastomer Actuators (DEA) are a low-TRL (Technology Readiness Level) technology that needs to prove its maturity and suitability to space implementation. In this work, the performances of a redundant manipulator based on DEAs are presented in terms of numerical and experimental results. A 4-DoF planar manipulator has been tested in a gravity-compensated setup. The system is composed by two double-cone actuators mounted in series, each of them providing actuation of two DoF. The end-effector is an optical marker whose position is detected by a vision system. The system has a total of four joint DoF and operates in the xy horizontal plane; only the x and y positions of the end-effector are controlled. Two degrees of redundancy are obtained and exploited for the optimization of joint torques to avoid the saturation of actuators. Numerical simulations have been conducted to predict the system behaviour. The laboratory facility emulates the zero-gravity orbital environment by means of a suspending cable. Detailed experimental results are presented and exploited for the validation of control algorithm and numerical models.

In this work, the mechanical and the self-healing behaviors of an ethylene-co-methacrylic acid ionomer were investigated in different testing conditions. The self-healing capability was explored by ballistic impact tests at low-velocity, midvelocity, and hypervelocity bullet speed; different experimental conditions such as sample thickness and bullet diameter were examined; in all impact tests, spherical projectiles were used. These experiments, in particular those at low and midspeed, allowed to define a critical ratio between sample thickness and bullet diameter below which full repair was not observed. After ballistic damage, the healing effi- ciency was evaluated by applying a pressure gradient through tested samples. Subsequently, morphology analysis of the affected areas was made observing all tested samples by scanning electron microscope. This analysis revealed different characteristic features of the damaged zones affected at different projectile speed. Stress-strain curves in uniaxial tension performed at different temperatures and strain rates revealed yield strength and postyield behavior significantly affected by these two parameters. A rise of temperature during high strain rate tests in the viscoplastic deformation region was also detected. This behavior has a strong influence on the self-repair- ing mechanism exhibited by the studied material during high-energy impact tests. V C 2013 Wiley Periodicals, Inc. J. Appl. Polym. Sci. 000: 000-000, 2013

In this study, multifunctional panels with structural and self-healing capabilities were experimentally investigated under high velocity impact conditions. The panels were composed of two layers, one made of fiber-reinforced plastics (FRP) and the other of a self-healing ionomer, assembled in four different configurations with surface density ranging from 6.85 to 7.49 kg/m2. A total number of 16 impact experiments are reported, where spherical aluminium projectiles with diameters from 2.3 to 4.5 mm were shot at impact velocity, which varied from 2.0 to 2.5 km/s. The panel response was analysed in terms of (1) impact damage ; and (2) capability to protect equipment located behind the panel from the cloud of fragments ejected downrange after panel perforation. Successful hole sealing was observed in panels consisting of a carbon FRP laminate and ionomer plate hit by projectiles up to 3.5 mm in diameter. The damage of the target was somewhat smaller when the ionomer was on the panel’s rear side. Such panels presented smaller debris cloud spread angle but lower momentum transfer to witness plates, suggesting that this panel configuration is preferable from the point of view of nearby equipment shielding effectiveness.

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 ° .

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.