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Heaterless hollow cathodes are a promising solution for reaching a good cathodic contact in electrodynamic tether systems. However, these cathodes cannot keep the plasma discharge if the electric current I C is below a threshold I m i n. If the tether is short, and to save power and expellant, an appropriate command strategy is needed to ensure that the cathode is triggered in conditions satisfying I C > I m i n . This study investigates four command strategies, including two fully autonomous approaches (one based on satellite position and the other on local plasma density) and two strategies involving ground segment access, differing in command generation location (ground segment and on-board computer). The success rates of the four strategies were studied though numerical simulation and for the specific scenario of the E.T.PACK demonstration mission. The results demonstrate that, as compared to continuous operation of the hollow cathode, the proposed strategies reduce propellant wastage by a factor between 5 and 10. The strategy requiring ground segment access and with the onboard computer making the command decision has the highest success rate (94%). Among the autonomous configurations, the strategy based solely on local plasma density achieves the highest success rate (69%). Additionally, the impact of sensor accuracy on success rates was assessed via Monte Carlo analysis. The command strategy relying on plasma density measurement is the most sensitivity strategy, with a 10% decrease in success rate due to sensor uncertainty. • Four strategies for commanding the hollow cathode of short electrodynamic tethers are proposed. • Duty cycle and rate of success are used to assess the performance of the strategies. • The two strategies involving contact with the ground segment have the best performance. • The two autonomous strategies only involve one onboard sensor and provide acceptable performance. [ABSTRACT FROM AUTHOR]

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. • Novel tether concept that combines a bare segment with thin-film photovoltaic cells. • Increased power generation enables consumable-less electrodynamic tether with high currents. • Attractive de-orbit technology which is less dependent on ambient conditions. [ABSTRACT FROM AUTHOR]

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 C 12 A 7 : 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. • The work plan of the E.T.PACK project is presented. • Goal: development of a deorbit kit based on a Low Work Function Tether (LWT). • The kit has 3 modes: as LWT and as standard bare tether in active and passive modes. • The performance of the kit is studied through numerical simulations. • The C12A7:e- electride is applied to four hardware elements. [ABSTRACT FROM AUTHOR]

Highlights • A constraint-free simulator for airborne wind energy systems is presented. • Fly- and ground-generation systems are considered. • Trade-off analysis of parallelization and lagrangian versus Hamiltonian formulations. • Self-consistent figure-of-eight trajectories are presented. • Open- and close-loop control strategies are considered. Abstract A parallelized flight simulator for the dynamic analysis of airborne wind energy (AWE) systems for ground- and fly-generation configurations is presented. The mechanical system comprises a kite or fixed-wing drone equipped with rotors and linked to the ground by a flexible tether. The time-dependent control vector of the simulator mimics real AWE systems and it includes the length of the main tether, the geometry of the bridle, the torque of the motor controllers of the rotors, and the deflections of ailerons, rudder and elevator. The use of a lagrangian formulation with a minimal coordinate approach and discretizing the main tether as a chain of inelastic straight rods linked by ideal (dissipative-less) rotational joints, yielded a non-stiff set of ordinary differential equations free of algebraic constraints. Several verification tests, including a reel-in maneuver that admits an analytical solution, are presented. The efficiency of the parallelization with the number of tether segments, and trade-off analysis of the lagrangian and hamiltonian formulations are also considered. The versatility of the simulator is highlighted by analyzing two maneuvers that are relevant for AWE scenarios. First, the simulator is used to compute periodic figure-of-eight trajectories with an open-loop control law that varies the geometry of the kite's bridle, as frequently done in ground-generation AWE systems. Second, an unstable equilibrium state of a tethered drone equipped with two rotors for energy harvesting is stabilized by implementing a closed-loop control strategy for the deflection of the control aerodynamic surfaces. [ABSTRACT FROM AUTHOR]

An analytical comparison of four technologies for deorbiting spacecraft from Low-Earth-Orbit at end of mission is presented. Basic formulas based on simple physical models of key figures of merit for each device are found. Active devices - rockets and electrical thrusters - and passive technologies - drag augmentation devices and electrodynamic tethers - are considered. A basic figure of merit is the deorbit device-to-spacecraft mass ratio, which is, in general, a function of environmental variables, technology development parameters and deorbit time. For typical state-of-the-art values, equal deorbit time, middle inclination and initial altitude of 850 km, the analysis indicates that tethers are about one and two orders of magnitude lighter than active technologies and drag augmentation devices, respectively; a tether needs a few percent mass-ratio for a deorbit time of a couple of weeks. For high inclination, the performance drop of the tether system is moderate: mass ratio and deorbit time increase by factors of 2 and 4, respectively. Besides collision risk with other spacecraft and system mass considerations, such as main driving factors for deorbit space technologies, the analysis addresses other important constraints, like deorbit time, system scalability, manoeuver capability, reliability, simplicity, attitude control requirement, and re-entry and multi-mission capability (deorbit and re-boost) issues. The requirements and constraints are used to make a critical assessment of the four technologies as functions of spacecraft mass and initial orbit (altitude and inclination). Emphasis is placed on electrodynamic tethers, including the latest advances attained in the FP7/Space project BETs. The superiority of tape tethers as compared to round and multi-line tethers in terms of deorbit mission performance is highlighted, as well as the importance of an optimal geometry selection, i.e. tape length, width, and thickness, as function of spacecraft mass and initial orbit. Tether system configuration, deployment and dynamical issues, including a simple passive way to mitigate the well-known dynamical instability of electrodynamic tethers, are also discussed. [ABSTRACT FROM AUTHOR]

A mathematical model of a kite connected to the ground by two straight tethers of varying lengths is presented and used to study the traction force generated by kites flying in cross-wind conditions. The equations of motion are obtained by using a Lagrangian formulation, which yields a low-order system of ordinary differential equations free of constraint forces. Two parameters are chosen for the analysis. The first parameter is the wind velocity. The second parameter is one of the stability derivatives of the aerodynamic model: the roll response to the sideslip angle, known also as effective dihedral. This parameter affects significantly the lateral dynamics of the kite. It has been found that when the effective dihedral is below a certain threshold, the kite follows stable periodic trajectories, and naturally flies in cross-wind conditions while generating a high tension along both tethers. This result indicates that kite-based propulsion systems could operate without controlling tether lengths if kite design, including the dihedral and sweep angles, is done appropriately. If both tether lengths are varied out-of-phase and periodically, then kite dynamics can be very complex. The trajectories are chaotic and intermittent for values of the effective dihedral below a certain negative threshold. It is found that tether tensions can be very similar with and without tether length modulation if the parameters of the model are well-chosen. The use of the model for pure traction applications of kites is discussed. [ABSTRACT FROM AUTHOR]

This study analyzes the performance of a tape-like bare electrodynamic tether as a promising propellant-free technology for the International Space Station (ISS) station-keeping, supporting the concept that the technology can provide significant mission benefits by reducing the ISS reliance on costly refueling operations for orbit maintenance. Convenient control laws for managing the electrical power supplied to the tether are proposed, exploring two distinct scenarios. The first involves using the electrodynamic tether continuously to counteract aerodynamic drag. The second adopts a cyclic approach, alternating between boosting the station with the tether and allowing for periods of natural decay. Optimal tether geometry, aimed at maximizing system efficiency, is also detailed. The study specifies an electrodynamic tether configuration featuring a 6-kilometer-long aluminum ribbon, 5 cm wide and 50 μ m thick, capable of overcoming aerodynamic drag ranging from 0.40 N to 0.80 N. Additionally, numerical simulations assess the tether performance under real environmental conditions. Furthermore, the study briefly introduces the potential of a photovoltaic tether as a fully autonomous system capable of supplying the necessary input power. • Electrodynamic tethers are proposed as a promising technology for ISS station-keeping • Air drag compensation and zig-zag are novel control strategies for orbit maintenance • Optimization codes are implemented for the identification of tether characteristics • The performances of the tether are investigated through numerical simulations • Bare Photovoltaic Tethers are proposed as autonomous systems for power harvesting [ABSTRACT FROM AUTHOR]

A low work-function tether (LWT), a subclass of bare electrodynamic tether made of a conductor partially coated with a low work function material, can exchange momentum and energy with planetary magnetospheres without any consumable. If fed by an onboard power source to reverse the natural direction of the current given by the motional electric field, a LWT can produce an useful thrust in drag compensation and re-boost scenarios. In the considered scheme, the LWT has an anodic bare segment for passive electron collection, followed by an insulated segment, a power source, and a cathodic segment that emits electrons passively through thermionic and photoelectric effects. Current and voltage profiles along the LWT are obtained and used to compute the system efficiency, i.e. the electrical power to mechanical power conversion rate, and the minimum electrical power to avoid space-charge effects in the cathodic segment. The design conditions to reach a high-efficiency regime are presented. For a given orbit and required thrust, optimal values for the tether geometry, including the fractional lengths of all three tether segments, are found. The results are applied to the preliminary design of a LWT system that would compensate the aerodynamic drag on the International Space Station. The analysis shows that the mission can be performed by a LWT fed with a power source of 4.7 kW, length, width and thickness equal to 5.1 km, 2 cm and 30 μ m , work function 1.5 eV, and temperature 600 K. [ABSTRACT FROM AUTHOR]

The increasing number of man-made objects in near-earth space is becoming a serious problem for future space missions around the Earth. Among the proposed strategies to face this issue, and due to the passive and propellant-less character, electrodynamic tethers appear to be a promising option for spacecraft in low Earth orbits thanks to the limited storage mass and the minimum interface requirements to the host spacecraft. This work presents the roadmap that the Electrodynamic Tether Technology for Passive Consumable-less Deorbit Kit (E.T.PACK) is following to develop a prototype of a deorbit device based on electrodynamic tether technology with Technology Readiness Level 4 by the end of 2022. The paper illustrates the roadmap of the activities carried out at the University of Padova, where software and hardware have been prepared to validate some of the critical elements of the deorbit device. Specifically, the software tools include: (a) the software called "DEPLOY" that allows the computation of a reference trajectory for the deployment of the tether and the completion of sensitivity analysis of the deployment trajectory to key error sources; (b) the software called "FLEXSIM" that predicts the performances of electrodynamic tethers as a function of the system configuration employed; and (c) the software called "FLEX" that includes the dynamical effects of tether flexibility and provides important information on the dynamic stability of the system during deployment and deorbiting phase. The paper describes in detail the three software tools and provides results of a simulation showing how it is possible to deorbit a 24-kg satellite from an initial orbital altitude of 600 km in less than 100 days using a 500-m long tape-like bare tether. The team has also developed laboratory mock-ups and performed experimental activities to: (a) determine the tether mechanical properties; (b) test the functionality of mechanisms used to deploy the tether; (c) test the functionality of the attitude control assembly used during the deployment phase; and (d) validate a passive damper designed for dissipating the longitudinal oscillations of the tether and thus guarantee the stability of the system during both deployment and deorbiting phase. The paper provides a description of both the laboratory setup and the experimental activities performed to validate EDT technologies, including the damping capability of a compact passive-damping mechanism, showing how it can reduce consistently the peak forces up to about 80%. • Optimization of full-length deployment trajectory. • High fidelity simulations of deorbiting with electrodynamic tethers. • Experimental validation of technologies for tape electrodynamic tethers. • Close-proximity deployment testing with frictionless table and trajectory control. • Development and testing of passive in-line damper. [ABSTRACT FROM AUTHOR]

De-orbiting satellites at end of mission would prevent generation of new space debris. A proposed de-orbit technology involves a bare conductive tape-tether, which uses neither propellant nor power supply while generating power for on-board use during de-orbiting. The present work shows how to select tape dimensions for a generic mission so as to satisfy requirements of very small tether-to-satellite mass ratio m t / M S and probability N f of tether cut by small debris, while keeping de-orbit time t f short and product t f × tether length low to reduce maneuvers in avoiding collisions with large debris. Design is here discussed for particular missions (initial orbit of 720 km altitude and 63° and 92° inclinations, and 3 disparate M S values, 37.5, 375, and 3750 kg), proving it scalable. At mid-inclination and a mass-ratio of a few percent, de-orbit time takes about 2 weeks and N f is a small fraction of 1%, with tape dimensions ranging from 1 to 6 cm, 10 to 54 μ m, and 2.8 to 8.6 km. Performance drop from middle to high inclination proved moderate: if allowing for twice as large m t / M S , increases are reduced to a factor of 4 in t f and a slight one in N f , except for multi-ton satellites, somewhat more requiring because efficient orbital-motion-limited electron collection restricts tape-width values, resulting in tape length (slightly) increasing too. [ABSTRACT FROM AUTHOR]