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Space missions in the environments of the outer planets require a large amount of propellant or assistant flybys. Moreover, the quadratic decrease of the radiation intensity with the distance from the sun makes solar panels less convenient as power source. Electrodynamic tether (EDT) technology is receiving increasing attention since it allows for obtaining both power and propulsion profiting from the high magnetic field and plasma density in the proximity of the giant planets. In particular, it has been shown that a self-propelled vehicle with a single EDT could be captured by Jupiter, were it not for the relativistic effects in the electron collection and the overheating of the tether due to the interaction with the environment. Here, we show that a multi EDT (MEDT) propulsion system can be used to circumvent these limitations. The suitable number, shape, and length of the tethers, depending on the size of the vehicle, are selected by structural and thermal analysis. A possible MEDT-propelled mission to Jupiter is described, including the design of the interplanetary trajectory, the orbital insertion, and the subsequent orbits around the planet. The results obtained demonstrate that such a bare MEDT system is a feasible option to provide propulsion and power in outer planets environments.

State-of-the-art electrodynamic tethers reach a steady electric current by using a bare segment to capture electrons passively from the ambient plasma (anodic contact) and an active electron emitter or a tether segment coated with a low-work-function material (cathodic contact) to emit electrons back and close the electrical circuit. This work proposes to take advantage of recent developments on thin-film solar cells and insert a photovoltaic (pv) tether segment in between the anodic and the cathodic contacts. Since thin-film solar cells can be folded and manufactured with any desired length and the same cross-section dimensions as the bare segment, i.e. width and thickness around few centimeters and tens of microns, the resulting device is compact and preserves bare tether simplicity. Detailed analysis of the current and voltage profiles throughout the tether shows that the electrical power introduced by the pv-segment into the tether-plasma circuit improves the performance and makes them less dependent on ambient conditions. The pv-segment decreases considerably the tether-to-plasma bias at the cathodic contact, thus opening the possibility to emit substantial current while using consumable-less electron emitters like thermionic and electron field emitters. The pv-segment also favors the current collection by increasing the tether-to-plasma bias at the bare segment. Propulsion and power generation applications and alternative architectures of bare-pv tethers are briefly discussed. This work received funding from the European Union’s Horizon 2020 research and innovation programme under grant agreement No 828902 (E.T.PACK project). GSA work is supported by the Ministerio de Ciencia, Innovación y Universidades of Spain under the Grant RYC-2014-15357.

The space debris occupies the orbit resources greatly, which seriously threats the safety of spacecraft for its high risks of collisions. Many theories about space debris removal have been put forward in recent years. The Electro Dynamic Tether (EDT), which can be deployed under gravity gradient, is considered to be an effective method to remove debris in low orbit for its low power consumption. However, in order to generate sufficient Lorentz force, the EDT needs to be deployed to several kilometers, which increases the risks of tangling and the instability of the EDT system. In the deployment process, different initial in-plane/out-of-plane angles, caused by direction error at initial release or the initial selection of ejection, affect the motion of EDT system seriously. In order to solve these problems, firstly, this paper establishes the dynamic model of the EDT system. Then, based on the model, safety metrics of avoiding tangling and assessing system stability during EDT deployment stage are designed to quantitatively evaluate the EDT system security. Finally, several numerical simulations are established to determine the safety ranges of the initial in-plane/out-of-plane angles on the EDT deployment.

Electrodynamic tethers can be employed as an immensely promising propulsion system for boosting the altitude of spacecraft, cite instance space stations, satellites, and the like, which ameliorates the expenditure of fuel and minimizes detrimental spacecraft impacts. This paper analyzes the libration dynamics and stability of reboost spacecraft with insulation electrodynamic tethers. A key aspect of spacecraft reboost with an electrodynamic tether is how to keep the tether aligned with the local vertical and stabilized in the context of external disturbances. It has been shown in the current research that the librational instability results in the tether slackness and swings in long-term motion without effective control. Moreover, electrodynamic force (Ampere force) is regarded as a distributed load acting on the tether causing tether deformations which may be detrimental if severe. Aiming at the problem, an optimal control method and PID are given to stabilize the libration motion by modulating the tether current. The dynamical model of the electrodynamic tether system is established using Lagrange equations of the second kind under considering tether deformation and control laws are proposed based on the model. The effectiveness of libration stability control is validated through numerical results in which a current regulation law with appropriate control parameters is used for the libration motion of electrodynamic tether system.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-03-10, 資料番号: PA2010045000

Electrodynamic tethers are proposed as propulsion and energy harvesters for space probes orbiting planets with a magnetic field and ionosphere, however there are no descriptions in the technical literature of the design of an electrical power system for such an application at subsystem and circuit detail. This paper presents a proposal for such a power system that extracts energy from the kinetic energy of a spacecraft using a bare electrodynamic tether, i.e. an unsheathed conductive wire or band, in low-Earth orbit. The application of the system is the powering of the spacecraft while in its final de-orbiting maneuvers at end-of-life with no reliance on the main spacecraft bus.

Proceeding of: Sixth International Conference on Tethers in Space, Madrid, June 12-14, 2019 The Electrodynamic Tether Technology for Passive Consumable-less Deorbit Kit (E.T.PACK) is a project aimed at the development of a deorbit kit based on low-work-function Tether (LWT) technology, i.e., a fully passive and electrically floating system made of a long conductive tape coated with a low-work-function material. The LWT interacts passively with the environment (ambient plasma, magnetic field, and solar radiation) to exchange momentum with the planet's magnetosphere, thus enabling the spacecraft to de-orbit and/or re-boost without the need for consumables. The main goal is to develop a deorbit kit and related software with Technology Readiness Level 4 and promote a follow-up project to carry out an in-orbit experiment. The planned kit in the experiment has three modes of operation: fully passive LWT and conventional electrodynamic tether equipped with an active electron emitter in passive and active modes. Several activities of the project pivot around the C12A7 : e- electride, which will be used in four hardware elements: (i) LWT (ii) hollow cathode, (iii) photo-enhanced thermionic emission device to convert solar photon energy into electrical energy, and (iv) a hollow cathode thruster. These elements, some of which do not belong to the deorbit kit, are synergetic with the main stream of the project and common to some tether applications like in-orbit propulsion and energy generation. This work explains the activities of E.T.PACK and the approach for solving its technological challenges. After reviewing past progresses on electrodynamic tethers and thermionic materials, we present a preliminary concept of the kit for the in-orbit experiment, some simulation results, and the key hardware elements. This work has received funding from the European Union's Horizon 2020 research and innovation programme under grant agreement No 828902 (E.T.PACK project). GSA work is supported by the Ministerio de Ciencia, Innovaci on y Universidades of Spain under the Grant RYC-2014-15357. SN work is supported by Comunidad de Madrid (Spain) under the Grant 2018/T2IND/11352

This paper investigates spacecraft orbital boost maneuver by a fully insulated electrodynamic tether using a coupled multiphysics dynamic model that considers electric-magnetic-dynamic energy coupling. The current-potential characteristics of the electrodynamic tether is described by a simplified analytical model, where revised Parker-Murphy model is employed to evaluate electric current generation at plasma contactors together with libration dynamics of tether. The 13th-order Earth's magnetic field model is used to account influence of inhomogeneity of Earth's magnetic field on orbital parameters of electrodynamic tether system, especially in inclined orbits. Parametric analysis has been conducted for orbital boost maneuver from 400 km to 1,200 km at four different orbital inclinations. The results show that (i) the orbit of electrodynamic tether system will change from circular to elliptical orbits, especially in inclined orbits, (ii) the voltage of onboard power supply at the anode affect electric current generation, (iii) the coupled multiphysics model is necessary to characterize the interaction of electrodynamic tether with the surrounding space environment and the controllability of electric current in electrodynamic tether, (iv) the motion of orbital boost maneuver by electrodynamic tether is stable, which is different from the deorbit by electrodynamic tether.

The paper considers parametric optimization of the process of changing orbital parameters by using a spinning electrodynamic tether system. Changes in the semi-major axis and eccentricity are taken as the two major goals, and two control laws are proposed accordingly. Current is regulated according to the instantaneous position of the conductive tether, which allows ensuring the calculated direction of the Lorentz force produced by the interaction of the conductive tether with the Earths magnetic field. A combined control scheme for simultaneous changes in the semi-major axis and eccentricity is proposed. The parameters of control laws are optimized on the basis of the Nelder-Mead method by using different objective functions and constraints. It is also shown that, by using the criteria of quick response and minimum impulse, we obtain optimal solutions corresponding to the boundary values of the selected parameters. Therefore, a convolution of these criteria is proposed as a compromise, which ensures a specified change in the orbital parameters of the system mass center.

形態: カラー図版あり, Physical characteristics: Original contains color illustrations, Accepted: 2020-01-20, 資料番号: PA2010046000

The dynamics analysis and libration suppression of an electrodynamic tether in circular orbits for a large spacecraft reboost are studied. The system is described by a dumbbell model that consists of two end-bodies, which are connected by one straight, massive and inextensible tether. A hybrid control methodology for the motion of electrodynamic tether system is synthesized to diminish the libration motions of the system. The tether current and thrust moment are used and regulated to remain stability of the libration motions and to boost the altitude of orbit continuously. Bellman’s dynamic programming is adopted to firstly develop linear regulators in linearized model, which are capable of damping out the in-planar and out-of-planar motions of the system, the regulators are then applied to the nonlinear system. The optimal control and quasi-optimal control are used for eliminating out-of-plane motion under different orbital inclinations. The results indicate that the quasi-optimal control methodologies have good performance in energy saving of total impulse. The results of simulation are shown to verify the successful of the devised control laws.

In this manuscript, a method for maneuvering a spacecraft using electrically charged tethers is explored. The spacecraft’s velocity vector can be modified by interacting with Earth’s magnetic field. Through this method, a spacecraft can maintain an orbit indefinitely by reboosting without the constraint of limited propellant. The spacecraft-tether system dynamics in low Earth orbit are simulated to evaluate the effects of Lorentz force and torques on translational motion. With 500-meter tethers charged with a 1-amp current, a 100-kg spacecraft can gain 250 m of altitude in one orbit. By evaluating the combined effects of Lorenz force and the coupled effects of Lorentz torque propagation through Euler’s moment equation and Newton’s translational motion equations, the simulated spacecraft-tether system can orbit indefinitely at altitudes as low as 275 km. Through a rare evaluation of the nonlinear coupling of the six differential equations of motion, the one finding is that an electrodynamic tether can be used to maintain a spacecraft’s orbit height indefinitely for very low Earth orbits. However, the reboost maneuver is inefficient for high inclination orbits and has high electrical power requirement. To overcome greater aerodynamic drag at lower altitudes, longer tethers with higher power draw are required.

In the last several years, the need for an alternative to chemical propulsive systems for low-orbit satellite deorbiting has become increasingly evident; a Tethered System can provide adequate thrust or drag without the complications of combustions and with a minimal impact on the environment. In this context, the authors are part of a team that is studying various tether applications and building a prototype of an electrodynamic tether system. The goal of this paper is to characterize tether materials in order to find valid solutions for future space tether missions. Mission requirements (e.g., the survivability to hypervelocity impacts and the capability to damp oscillations in electrodynamic tethers) influence the choice of tether parameters such as cross section geometry (round wires or tapes), materials, length, and cross section sizes. The determination of the elastic characteristics and damping coefficients is carried out through a campaign of experiments conducted with both direct stress/strain measurements and the laboratory facility SPAcecRraft Testbed for Autonomous proximity operatioNs experimentS (SPARTANS) on a low friction table at the University of Padova. In the latter case, the stiffness and damping of a flexible line were verified by applying different tensile load profiles and then measuring the tether-line dynamic response in terms of tension spike amplitude, oscillation decay, and estimation of the damping coefficient.

Algorithms for stabilizing the motion of an electrodynamic tether system in the orbit of an Earth satellite are analyzed. The system under consideration consists of two small spacecraft and a tether connecting them. A tether system is designed to change its orbital parameters, using the interaction of a current-conducting tether with the Earth’s magnetic field. Several mathematical models are used to describe the movement of the system, with varying degrees of detail describing its movement. An algorithm for stabilizing the movement of the tether system in the vicinity of the local vertical is proposed. A feature of the considered algorithm, based on the feedback principle, is the stabilization of the bending vibrations of the tether from the action of the distributed load acting on the current-conducting tether in a magnetic field. It is shown that not taking into account the bending vibrations of the tether in the stabilization algorithms can lead not only to a deterioration in the quality of stabilization, but also to a loss of stability of the system motion (its transition to rotation).

The work done on probability of collision between spherical objects in orbit is extended here to the case of one spherical object and one circular or rectangular object. The former is a model for s...

This paper considers the motion of a low-orbit electrodynamic tether system designed to raise the orbit of a small spacecraft or nanosatellites. The system operates in the thrust generation mode. The orbit is raised by the Ampere force resulting from the interaction of the conducting tether with the Earth’s magnetic field. The mathematical model of motion is constructed using the Lagrange method taking into account the effect of the distributed loads from the Ampere force and the aerodynamic forces on the tether. It is shown that the motion of the system relative to the center of mass is unstable if the current is constant. It is proposed to use a linear regulator to stabilize the motion of the system with respect to the local vertical. Bellman’s dynamic programming principle is used to synthesize the regulator.

Electrodynamic tether technology is a novel way to perform spacecraft orbital transfers. This paper focuses on a bare electrodynamic tether (BEDT) system and evaluates its orbital transfer performa...

The electric sail is an innovative propellantless propulsion concept that gains continuous momentum from the solar wind. In this work, the electric sail’s attitude dynamics is established from a mu...

This paper aims to highlight the attitude determination and control of a 2-U class Nanosatellite in a decaying orbit. The ever-increasing space debris in the Low Earth Orbit (LEO) poses a significant threat to operational spacecraft. This paper aims to describe an inexpensive, space and mass efficient passive deorbiting system using an electrodynamic tether. It works as an independent and autonomous system, using only electrodynamic drag to achieve deorbit. An attitude control system to ensure the working of payloads onboard the satellite and establish communication with the ground station during deorbiting has been further discussed. The attitude of the satellite is determined using onboard sensors: Sun sensors, a three-axis magnetometer, and a gyroscope. An orbit propagator (OP) model has been used to estimate the orbital parameters calibrated once per orbit utilising a GPS. A Quaternion Estimator (QuEst) and proportional-derivative (PD) controller is used for calculating the desired rotation and the corresponding torque required to orient the satellite. A complete three-axis stabilisation is achieved using the two actuators onboard the satellites: the magnetorquers and three-axis reaction wheels that apply the desired torque. The voltage developed in the tether system and accordingly the force produced have been modelled. The results indicate a period of less than one year for complete deorbiting. An extensive analysis of the control of the satellite and the force acting on the satellite-tether system during the ejection phase and a deorbiting phase has been discussed, and the simulation results have been provided.

Spinning electrodynamic tether systems are considered ideal platforms for payload transportation, removal of space debris, artificial gravity, and so on, for they provide a propellantless s...

Power generation can be realized in space when current is induced on a bare electrodynamic tether system. The performance of power generation is discussed based on a debris mitigation mission by numerical simulation in the paper. A Li-ion battery subsystem is used to complete the energy conversion—harvest and supply the energy. The battery can provide 10–300 W average electric power continuously during several hundred hour mission time. The energy conversion efficiency ranges from 1% to a maximum value 30%. With constant power consumption on board, the battery operation generally experiences a discharging phase, a charging phase, and a stable phase. The first two phases determine the mission risk coefficient. The heating problem in the stable phase cannot be ignored. The optimization of battery design and tether design should be considered for each debris mitigation mission. An extra control circuit or small battery voltage with large capacity for battery design is suggested to eliminate the stable phase. Wide or long tether designs are more appropriate for mission with high or low power demands on board, respectively. The power generation is affected by the system mass and the mission orbit parameters.

The number of space debris orbiting our Earth has been continuously increasing since the beginning of the space era. The space community is converging on responsible conducts and self-regulations to address this serious problem that is degrading the near-Earth environment. In this context, green deorbiting technologies and strategies alternative to the traditional chemical propulsion are under investigation, including Electrodynamic Tethers (EDTs) because they are a promising option. To increase EDT technology maturity level, some critical points shall be addressed and experimentally evaluated, including the deployment of tape tethers, to demonstrate their reliability. This paper presents results of an experimental validation of the Deployment Mechanism (DM) proposed for the H2020 FET OPEN Project E.T.PACK - Electrodynamic Tether Technology for Passive Consumable-less Deorbit Kit. We developed a mockup that hosts the DM and other elements that are on board the tip mass of a tethered system, using off-the-shelf components. The DM is tested for the first part of the tether deployment maneuver employing the SPARTANS facility of the University of Padova. This facility includes a Testing Table where the mock-up can move with almost no friction and a Motion Capture system that provides an accurate estimation of the mock-up motion during this first part of the tether deployment maneuver.

Due to its extraordinary properties, the C12A7:e− electride has been suggested in the past as one of the most promising materials for coating Low Work-function Tethers (LWTs). Such subclass of electrodynamic tether, made of a conductive substrate coated with materials that enhance the electron emission through the thermionic and photoelectric effects, would constitute a fully passive and consumable-free device for deorbiting space debris from Low Earth Orbit. This work presents experimental results on the manufacturing and testing of small tape-like LWT samples, which is the relevant geometry for electrodynamic tether applications. After preparing C12A7:e− targets under specific and controlled conditions, thin titanium foils have been coated with the C12A7:e− electride by using two different physical vapor deposition techniques: magnetron sputtering (MS) and pulsed laser deposition (PLD). In the case of MS, important difficulties and defects were found, including target damage, poor growth rate, and oxidation and changes in the composition of the coating. However, the performance on the PLD coating was radically different. First, Rutherford backscattering spectrometry confirmed that the composition of the thin film with the PLD coincides with that of the target. Second, X-ray photoelectron spectroscopy and thermionic emission experiments showed that the work function of the LWT sample is 2.6 ± 0.1 eV. Therefore, the trade-off analysis indicates that PLD can be used to manufacture LWT samples with work function close to the one of the bulk C12A7:e− (2.4 eV). Nevertheless, its application as passive cathode in electrodynamic tethers still requires further research activities. This work was supported by Agencia Estatal de Investigación (Ministerio de Ciencia, Innovación y Universidades of Spain) under the project ESP2017-82092-ERC (AEI). Work by JFP and AP has been supported by the Centro para el Desarrollo Tecnológico e Industrial (CDTI) under the project IDI-20171255. FJP and PT acknowledge the support by ATD under project contract ICMM-CSIC #010101170104. SN work is supported by Comunidad de Madrid (Spain) under the Grant 2018/T2IND/11352. GSA work is supported by the Ministerio de Ciencia, Innovación y Universidades of Spain under the Grant RYC-2014-15357.

This paper is devoted to the problem of a small payload delivery from an orbit to the Earth by a reusable space tether system consisting of a satellite in circular orbit, a tether, and a bottom mod...

Various promising applications of electrodynamic tether have been proposed for space missions over the past decades. A crucial issue of these missions is to deploy an electrodynamic tether under a rapid and stable state. This paper aims to stabilize the libration motions of a bare electrodynamic tether during its three-dimensional deployment. The tethered system under consideration consists of a main-satellite and a sub-satellite connected to each other through a bare electrodynamic tether. A widely used dumbbell assumption considering the tether as rigid and inflexible is adopted to facilitate the dynamic modeling and analysis of the tethered system. A pair of active control laws is synthesized by simultaneously regulating the electric current and tether tension to achieve an efficient stabilization of the three-dimensional libration of the bare electrodynamic tether in the deployment process. Moreover, comparisons of three groups of numerical simulations are performed to evaluate the in uences of orbital inclinations and geomagnetic field models and the performance of the active control laws.

This paper provides a structural evaluation for electrodynamic tape tethers that can resist collisions with small items of space debris. To actively remove space debris such as defunct satellites f...

This paper investigates the applicability of five active methods for removal of large low Earth orbit debris. The removal methods, namely net, laser, electrodynamic tether, ion beam shepherd, and r ...

This paper studies the control of libration stability of an insulated electrodynamic tether system in orbital boost. The electrodynamic tether is assumed rigid and the tethered spacecraft are modeled as lumped masses, while a complete electric circuit model is established by the Parker-Murphy model and Kirchhoff's first law. The electric current regulation is realized by adjusting the resistance of a resistor in the circuit. The impact of periodic variations of electrodynamic force on the libration of electrodynamic tether is minimized by a simple fuzzy-based continuous electric current control. The proposed controller maintains the merits of continuous electric current control and simplicity of on-off current control. The effectiveness of the fuzzy-based controller has been demonstrated by analyzing the libration dynamics of electrodynamic tether with continuous and discrete on-off current profiles in orbital boost. Numerical results show that the proposed approach is much effective than the on-off current control in stabilizing both in-plane and out-of-plane libration of the electrodynamic tether system subjected to periodic and non-periodic perturbations.

A comprehensive tour of the complex outer planet systems is a central goal in space science. However, orbiting multiple moons of the same planet would be extremely prohibitive using traditional propulsion and power technologies. In this paper, a new mission concept, named Magnetour, is presented to facilitate the exploration of outer planet systems and address both power and propulsion challenges. This approach would enable a single spacecraft to orbit and travel between multiple moons of an outer planet, without significant propellant or onboard power source. To achieve this free-lunch ‘Grand Tour’, Magnetour exploits the unexplored combination of magnetic and multi-body gravitational fields of planetary systems, with a unique focus on using a bare electrodynamic tether for power and propulsion. Preliminary results indicate that the Magnetour concept is sound and is potentially highly promising at Jupiter.

An electrodynamic-tether system in low Earth orbit has the possibility of transferring energy from its orbital potential energy to its electrical system. Thus, electrical energy can be made availab...

The paper develops a coupled multiphysics finite element method to analyze satellite deorbit by a bare flexible electrodynamic tether. Unlike the existing approaches, the current method assumes that the tether is flexible and deflectable, and its efficiency of electron collection varies along the tether length depending on tether deflection. The orbital motion limited theory, which dictates the electron collection by a bare tether, is discretized and solved with the same finite element mesh as the tether dynamics to couple the electron collection with the tether flexural deflection. The advantages of the new method are demonstrated by numerical simulations. Compared with a reference method based on straight tether assumption, the coupling effect between the electron collection and tether deflection is significant, leading to the dynamic variation of electrodynamic force acting on the tether. Although the deorbit rates predicted by these two methods are almost the same, the new method predicts a shorter st...

The success of nanosatellites (1–10 kg) and the miniaturization of sophisticated low-power electronics has motivated interest in even smaller “smartphone”-sized spacecraft as either standalone spacecraft or elements in a maneuverable fleet. These spacecraft, known as picosatellites (100 g–1 kg) and femtosatellites (less than 100 g), have the potential to enable missions requiring a distributed fleet of sensor spacecraft (for example, distributed aperture, simultaneous spatial sampling, etc.). However, without some degree of propulsion capability, these spacecraft would behave more as an uncontrolled swarm than as a coordinated formation. Furthermore, lifetime in low Earth orbit can be limited for low-mass spacecraft with high area-to-mass ratios. This paper shows that a relatively short (few meters) electrodynamic tether is capable of providing picosatellites and femtosatellites with propellantless drag cancellation and even the ability to change orbit over an altitude range determined by the ionospheric ...

The problem of modeling the braking process of a nanosatellite using a current-conducting tether is considered. The task is related to the actual problem of removing nanosatellites from space that have exhausted their life without the use of jet engines. The interaction of the conducting current of the tether with the magnetic field of the Earth leads to the appearance of the ampere (Lorentz) force that provides braking of the system and a steeper trajectory of the nanosatellite entry into the dense layers of the atmosphere. Modeling and analysis of the braking system is carried out in two models: the model for a straight tether and the model with distributed parameters where the tether is represented by a set of material points. Comparison of the considered equations of motion of the system in terms of estimating the increments of the orbital parameters of the system is considered. The relative influence of aerodynamic forces and ampere forces on the braking process of the system is estimated by the method of mathematical modeling. As an example, a numerical simulation of an electrodynamic tether system motion with a bare tether in the earth's magnetic field is considered.

This paper aims to characterize the forces acted upon a 2U-class Nano-satellite by a passive electrodynamic tether system used as an inexpensive, space and mass efficient de-orbiting mechanism in the Low Earth Orbit (LEO). As stated by the IADC guidelines, an object in the low earth orbit should not have an orbital lifetime exceeding 25 years which can be attained by the satellites themselves due to significant atmospheric drag. The de-orbiting system involves a spool with an electrically conducting thread wound around it. It is ejected using a spring-loaded mechanism with the other end of the thread secured to the satellite body. The shape of the spool and the type of winding have been optimized to ensure maximum packing efficiency and a stable ejection. It has resulted in a tether length of around 300 meters for the considered system. The mass of the spool with the tether winded upon it is about 80 grams (mass of the tether being 32 grams) which makes it a suitable solution for small satellites with stringent mass constraints. Study of the forces acting on the satellite body during the deployment becomes essential to avoid the destabilization of the satellite which would result in the winding of the tether about the satellite body. Excessive tension on the tether during the deployment phase might result in a failure of the thread material. The forces acting on the system during the deployment phase have been analyzed. Methods to reduce and dampen these destructive effects have been discussed in detail. Once deployed, various forces act on the satellite system due to the interaction of the conductive thread with the time-varying magnetic field and the ambient plasma. Being a passive tether system, it experiences a change in the direction of the magnetic field while crossing over the poles resulting in the force to be applied in the same direction in different parts of the orbit. This force, always opposing the motion of the satellite, increases the eccentricity of the satellite orbit causing it to burn up at the perigee in under 2 years as opposed to 25 years. A detailed study on the respective forces and its effects on the satellite's orbit has been carried out in this paper.

This works presents a system level analysis of a Deorbit Kit (DK) based on electrodynamic tether technology. The analysis is focused on two relevant scenarios for deorbiting space debris: (i) Earth Observation (EO) satellites with mass in the range of 700 kg -1000 kg and initial orbital altitude of 800 km and 98° inclination, and (ii) Mega Constellation (MC) spacecraft in the order of 200 kg and initial orbit at 1200 km of altitude and 90° of inclination. The scenarios have been selected considering the orbits that are already suffering from the space debris problem or will suffer in the next future. The DK implements a bare electrodynamic tether for capturing electrons passively from the ambient plasma while three different methods are considered for emitting the electrons back to the plasma to reach a steady electrical current on the tether. The three studied options to close the electrical circuit are: (a) a hollow cathode, which has a high technological maturity but needs expellant and a little of power, (b) a thermionic emitter, which does not involve expellant but needs power, and (c) a Low Work-function Tether (LWT) that does not need neither expellant nor power because it has a segment coated with a special material that emits electrons passively through the thermionic and photoelectric effects. In order to provide a fully autonomous operation even in case of critical failure of the mother spacecraft, the DK includes a deployment mechanism, a telemetry and telecommand system, a complete Attitude Determination and Control System with attitude sensors (GNSS, sun sensors, magnetometer) and actuators (magneto torquers), solar panels and batteries. Upon activation, the DK autonomously de-tumbles the satellite, deploys a tether and carries out the satellite's de-orbiting. The study presents DK architectures, mass budgets and simulation results for the two scenarios. It is shown that a complete DK with mass below 6% the mass of the host spacecraft can deorbit EO and MC satellites in about 1.5 years and 10 years, respectively. The importance of the development of the LWT concept to enhance the simplicity and reduce the mass, power and volume budget is highlighted. This work was supported by the European Union's Horizon 2020 Research and Innovation Programme under grant agreement No.828902 (E.T.PACK project). GSA work is supported by the Ministerio de Ciencia, Innovación y Universidades of Spain under the Grant RYC-2014-15357.

A deployed electrodynamic tether, provided with a tip-mass, can be used to deorbit a spacecraft at the end of its operative life. During the deployment of the tether from the spacecraft, the distance measurement between the spacecraft and the tip-mass may be useful to control the whole system. In the present work, three vision-based approaches to measure the tip-mass distance and their achievable uncertainties are compared. The first approach employs two cameras (a stereo camera) to evaluate the tip-mass distance using triangulation. The second and the third approaches are based on the size of the tip-mass projection in the image plane of a monocular camera, respectively when the tip-mass is inside (in focus) or outside (out of focus) the depth of field of the vision system. The obtained results highlight the advantage of the stereo approach, but also the promising behavior of the approach based on defocus. The worst approach is always the monocular one based on the in-focus size of the projected tip-mass.

The use of a hollow cylindrical mesh for an electrodynamic tether to improve its survivability in space is proposed. The proposed tether is expected to be used in space missions to remove uncontrollable satellites in orbit. Although the mass of the hollow cylindrical tether is the same as that of a conventional solid and cylindrical (i.e., round) tether, it has a larger diameter, which improves its survivability. To investigate the damage to the proposed hollow cylindrical mesh tether by debris impact, hypervelocity-impact experiments were performed by using a two-stage light-gas gun. A high-speed camera was used to photograph the debris cloud generated by the impact of debris on the mesh tether, and the damaged area of the mesh tether after impact was compared with the expansion size of the debris cloud. Numerical simulations were conducted by using the smoothed-particle hydrodynamics method to interpolate the experimental data. The experimental and simulation results confirmed the potentially increased ...

Electrodynamic tethers are used to remove large uncontrollable satellites from low Earth orbits. However, tether systems are vulnerable to collisions with space debris orbiting at high speeds. Such collisions can easily sever electrodynamic tethers. The success of an electrodynamic tether mission to remove uncontrollable satellites from low Earth orbits will depend in part on how often the electrodynamic tether collides with space debris and how the electrodynamic tether fractures as a result of a collision. The dynamic fracture progress of an electrodynamic tether in response to a collision with space debris has not been previously considered in predicting the success rate of an electrodynamic tether mission. In this study, the fracture progress of an electrodynamic tether owing to a hypervelocity impact collision with space debris was experimentally and numerically examined to understand the dynamic fracture phenomenon. In addition, an estimate of the survivability of an electrodynamic tether in orbit w...

This paper focuses on the feasibility of constellation maintenance for large-scale satellite constellations (i.e., mega-constellations) by the electrodynamic tethered multi-agent system with consideration of electric-magnetic-dynamic energy coupling in the orbital maneuver process. The electrodynamic tethered multi-agent system consists of electrodynamic tethers and agents (satellites, robots, or spacecrafts). The nonlinear dynamic characteristics of the electrodynamic tethered multi-agent system are investigated here, to analyze the performance of orbital maneuver. The electrodynamic tether is divided into two parts, such that, the fully insulated tether and the fully bare tether, which used to boost and deboost the electrodynamic tethered multi-agent system respectively. In the orbital boost phase, the real-time electric current generated from the segment of the fully insulated tether. Similarly, in the orbital deboost phase, the electric current generated from the fully bare tether segment. The numerical results show that the system could accelerate or decelerate the motion of the electrodynamic tethered multi-agent system by regulating the magnitude and polarity of electric current carrying tether to generate a Lorentz force, leading to orbital boost or deboost. The electrodynamic tether propulsion is useful for the constellation maintenance, such that, the Starlink, LeoSat and Project Kuiper satellite constellations.

Spinning electrodynamic tether systems are good platforms for payload transportation, but require orbit reboosts after one or several tossing manoeuvres. In this paper an impact analysis is proposed to predict the change of system orbits after payload tossing. Changes of key parameters (eccentricity, semi-major axis and tether spinning rate) are derived. Based on previous analysis, three control schemes are proposed to regulate current to reboost system orbits back to its original ones. One mission is assumed in numerical parts to validate the effectiveness of the proposed impact analysis and control schemes.

A carbon nanotube-based field emission cathode (FEC) has been researched and developed by the Japan Aerospace Exploration Agency mainly for space debris removal systems with electrodynamic tether (EDT) propulsion. An inorbit experiment to demonstrate the EDT technologies including the FEC was planned and conducted in early 2017. The FEC operated well without any critical trouble during the one-week mission period. The data obtained includes its electron emission capability to space plasma and its durability in low Earth orbit, which suggest the prospect of using the FEC in space.

One of the key functionalities required in a mission for removing a noncooperative spacecraft is the assessment of its kinematics and inertial properties. In a few cases, this information can be ap...