Special Issue "Electric Propulsion"

A special issue of Aerospace (ISSN 2226-4310).

Deadline for manuscript submissions: 30 September 2020.

Special Issue Editor

Prof. Dr. Richard E. Wirz
Website
Guest Editor
Department of Mechanical and Aerospace Engineering, UCLA, Los Angeles, CA USA
Interests: electric propulsion; plasma; renewable energy; energy storage

Special Issue Information

Dear Colleagues,

Electric propulsion (EP) has led to a revolution in space mission capabilities, including unprecedented specific impulse, mission deltaV, thrust precision, and spacecraft control authority. These advancements are the result of significant experimental and modeling efforts, as well as technology demonstration missions. The space community is at an important stage with the development of new and exciting concepts for deep space and Earth-orbiting missions, including missions to the Moon, Mars, asteroids, and beyond; unprecedented space-based observatories; large satellites; small/nano satellites; satellite formations; and many other space missions—most of which can benefit from EP. This Special Issue on electric propulsion aims to discuss where we are, how we got here, and where we are headed with regard to current and emerging EP technology, experiments, modeling, theory, related science, and mission capabilities. Submissions are encouraged from all researchers engaged in EP who would like to be a part of this conversation.

Prof. Dr. Richard E. Wirz
Guest Editor

Manuscript Submission Information

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Published Papers (6 papers)

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Research

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Open AccessArticle
Characterization of Vacuum Arc Thruster Performance in Weak Magnetic Nozzle
Aerospace 2020, 7(6), 82; https://doi.org/10.3390/aerospace7060082 - 19 Jun 2020
Abstract
Vacuum arc thruster performance in a magnetic nozzle configuration is experimentally characterized. Measurements are performed on a miniature coaxial thruster with an anode inner diameter of 1.8 mm. The magnetic field B is produced by a single air coil, 18 mm in diameter. [...] Read more.
Vacuum arc thruster performance in a magnetic nozzle configuration is experimentally characterized. Measurements are performed on a miniature coaxial thruster with an anode inner diameter of 1.8 mm. The magnetic field B is produced by a single air coil, 18 mm in diameter. Direct measurement of thrust, mass consumption and arc current are performed. To obtain statistically viable results 6000 arc pulses are analyzed at each operational point. Cathode mass erosion is measured using laser profilometry. To sustain thruster operation over several measurement cycles, an active cathode feeding system is used. For 0 < B 0.2 T, performance increase over the non-magnetic case is observed with the best thrust to arc power ratio T / P 9 μ N/W obtained at B 0.2 T. A parametric model is provided that captures the performance enhancement based on beam collimation and acceleration by the magnetic nozzle. For B > 0.2 T, the arc discharge is shown to be suppressed nullifying any additional gains by the nozzle effect. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Open AccessArticle
Sequential Low-Thrust Orbit-Raising of All-Electric Satellites
Aerospace 2020, 7(6), 74; https://doi.org/10.3390/aerospace7060074 - 04 Jun 2020
Abstract
In this paper, we consider a recently developed formulation of the electric orbit-raising problem that utilizes a novel dynamic model and a sequence of optimal control sub-problems to yield fast and robust computations of low-thrust trajectories. This paper proposes two enhancements of the [...] Read more.
In this paper, we consider a recently developed formulation of the electric orbit-raising problem that utilizes a novel dynamic model and a sequence of optimal control sub-problems to yield fast and robust computations of low-thrust trajectories. This paper proposes two enhancements of the computational framework. First, we use thruster efficiency in order to determine the trajectory segments over which the spacecraft coasts. Second, we propose the use of a neural network to compute the solar array degradation in the Van Allen radiation belts. The neural network is trained on AP-9 data and SPENVIS in order to compute the associated power loss. The proposed methodology is demonstrated by considering transfers from different geosynchronous transfer orbits. Numerical simulations analyzing the effect of thruster efficiency and average power degradation indicate the suitability of starting the maneuver from super-geosynchronous transfer orbits in order to limit fuel expenditure and radiation damage. Furthermore, numerical simulations demonstrate that proposed enhancements are achieved with only marginal increase in computational runtime, thereby still facilitating rapid exploration of all-electric mission scenarios. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Open AccessArticle
Impulse and Performance Measurements of Electric Solid Propellant in a Laboratory Electrothermal Ablation-Fed Pulsed Plasma Thruster
Aerospace 2020, 7(6), 70; https://doi.org/10.3390/aerospace7060070 - 30 May 2020
Abstract
Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed [...] Read more.
Electric solid propellants are advanced solid chemical rocket propellants that can be controlled (ignited, throttled and extinguished) through the application and removal of an electric current. This behavior may enable the propellant to be used in multimode propulsion systems utilizing the ablative pulsed plasma thruster. The performance of an electric solid propellant operating in an electrothermal ablation-fed pulsed plasma thruster was investigated using an inverted pendulum micro-newton thrust stand. The impulse bit and specific impulse of the device using the electric solid propellant were measured for short-duration test runs of 100 pulses and longer-duration runs to end-of-life, at energy levels of 5, 10, 15 and 20 J. Also, the device was operated using the current state-of-the-art ablation-fed pulsed plasma thruster propellant, polytetrafluoroethylene (PTFE). Impulse bit measurements for PTFE indicate 100 ± 20 µN-s at an initial energy level of 5 J, which increases linearly with energy by approximately 30 µN-s/J. Within the error of the experiment, measurements of the impulse bit for the electric solid propellant are identical to PTFE. Specific impulse when operating on PTFE is calculated to be about 450 s. It is demonstrated that a surface layer in the hygroscopic electric solid propellant is rapidly ablated over the first few discharges of the device, which decreases the average specific impulse relative to the traditional polytetrafluoroethylene propellant. Correcting these data by subtracting the early discharge ablation mass loss measurements yields a corrected electric solid propellant specific impulse of approximately 300 s. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Open AccessArticle
Far-Field Plume Characterization of a 100-W Class Hall Thruster
Aerospace 2020, 7(5), 58; https://doi.org/10.3390/aerospace7050058 - 14 May 2020
Abstract
The 100 W-class ISCT100-v2 Hall Thruster (HT) has been characterized in terms of far-field plume properties. By means of a Faraday Cup and a Retarding Potential Analyzer, both the ion current density and the ion energy distribution function have been measured over a [...] Read more.
The 100 W-class ISCT100-v2 Hall Thruster (HT) has been characterized in terms of far-field plume properties. By means of a Faraday Cup and a Retarding Potential Analyzer, both the ion current density and the ion energy distribution function have been measured over a 180 circular arc for different operating points. Measurements are compared to far-field plume characterizations performed with higher power Hall thrusters. The ion current density profiles remain unchanged whatever the HT input power, although an asymptotic limit is observed in the core of the plume at high discharge voltages and anode mass flow rates. In like manner, the ion energy distribution functions reveal that most of the beam energy is concentrated in the core of the plume [ 40 ; 40 ] . Moreover, the fraction of low energy ion populations increases at large angles, owing to charge exchange and elastic collisions. Distinct plume regions are identified; they remain similar to the one described for high-power HTs. An efficiency analysis is also performed in terms of current utilization, mass utilization, and voltage utilization. The anode efficiency appears to be essentially affected by a low voltage utilization, the latter originating from the large surface-to-volume ratio inherent to low-power HTs. Experimental results also show that the background pressure clearly affects the plume structure and content. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Open AccessArticle
Modeling and Characterization of a Thermally Controlled Iodine Feeding System for Electric Propulsion Applications
Aerospace 2020, 7(2), 10; https://doi.org/10.3390/aerospace7020010 - 23 Jan 2020
Abstract
Iodine is considered as a feasible alternative to xenon as a propellant for electric propulsion systems, thanks to its good propulsive performance, high availability, and high storage density. However, as iodine is stored in solid state at ambient temperature, current state-of-the-art propellant management [...] Read more.
Iodine is considered as a feasible alternative to xenon as a propellant for electric propulsion systems, thanks to its good propulsive performance, high availability, and high storage density. However, as iodine is stored in solid state at ambient temperature, current state-of-the-art propellant management systems are not suitable to be used with it. Moreover, due to its high reactivity, iodine imposes requirements on material-compatibility, hindering the use of mass flow measurement and control systems typically used with other propellants. The architecture of a controlled iodine feeding system for low power (200 W class) ion and Hall effect thrusters is presented and the resulting prototype is described. It consists of a sublimation assembly whose temperature is used to control the tank pressure, a normally-closed ON-OFF valve, and a thermal throttle to perform the fine control of the mass flow rate. A 1D thermal-fluid model concerning the vapor generation in the tank, and its evolution along the different components is detailed. The thermal throttle model has been experimentally verified using air as a working fluid. The model results agree with the measurements of the verification tests in the hypothesis of the presence of an extended region at the entrance of the pipe where the laminar flow velocity and temperature profiles are not fully developed (known as entry flow region). Finally, the system is experimentally characterized and the model of the full system is calibrated using experimental measurements. The calibration shows that the thermal throttle flow presents an entry flow region, that the viscosity is correctly modeled, and that there is a difference between the measured tank temperature and the effective sublimation temperature. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Review

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Open AccessReview
A Review of Low-Power Electric Propulsion Research at the Space Propulsion Centre Singapore
Aerospace 2020, 7(6), 67; https://doi.org/10.3390/aerospace7060067 - 28 May 2020
Abstract
The age of space electric propulsion arrived and found the space exploration endeavors at a paradigm shift in the context of new space. Mega-constellations of small satellites on low-Earth orbit (LEO) are proposed by many emerging commercial actors. Naturally, the boom in the [...] Read more.
The age of space electric propulsion arrived and found the space exploration endeavors at a paradigm shift in the context of new space. Mega-constellations of small satellites on low-Earth orbit (LEO) are proposed by many emerging commercial actors. Naturally, the boom in the small satellite market drives the necessity of propulsion systems that are both power and fuel efficient and accommodate small form-factors. Most of the existing electric propulsion technologies have reached the maturity level and can be the prime choices to enable mission versatility for small satellite platforms in Earth orbit and beyond. At the Plasma Sources and Applications Centre/Space Propulsion Centre (PSAC/SPC) Singapore, a continuous effort was dedicated to the development of low-power electric propulsion systems that can meet the small satellites market requirements. This review presents the recent progress in the field of electric propulsion at PSAC/SPC Singapore, from Hall thrusters and thermionic cathodes research to more ambitious devices such as the rotamak-like plasma thruster. On top of that, a review of the existing vacuum facilities and plasma diagnostics used for electric propulsion testing and characterization is included in the present research. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Planned Papers

The below list represents only planned manuscripts. Some of these manuscripts have not been received by the Editorial Office yet. Papers submitted to MDPI journals are subject to peer-review.

Title: editorial
Authors: Igor LEVCHENKO
Affiliation: Plasma Sources and Applications Centre | Space Propulsion Centre, Singapore NIE, Nanyang Technological University

Title: Cutting-Edge Arc-based Plasma Propulsion
Authors: G Herdrich
Affiliation: Priv. Doz. Dr.-Ing. Georg Herdrich Adj. Assoc. Prof./ Baylor University Institute of Space Systems Head Plasma Wind Tunnels and Electric Space Propulsion Pfaffenwaldring 29 70569 Stuttgart Germany

Title: Selected Advanced Plasma Propulsion Systems
Authors: G Herdrich
Affiliation: Priv. Doz. Dr.-Ing. Georg Herdrich Adj. Assoc. Prof./ Baylor University Institut für Raumfahrtsysteme Leiter Plasmawindkanäle und Elektrische Raumfahrtantriebe Pfaffenwaldring 29 70569 Stuttgart

Title: A Future Prospect of Electric Propulsion for Solar System Explorations
Authors: Ikkoh Funaki
Affiliation: Japan Aerospace Exploration Agency
Abstract: Future prospect toward solar system exploration is discussed based on a brief overview of electric propulsion (EP) activities and an estimation of future activities. Currently, EP based projects for Lunar gateway and several probes for small bodies such as asteroids are going on, and higher energy missions such as Mars sample return is under planning. The trend required for EP based interplanetary spacecraft is shifting toward higher energy either in a short term (several years) or in a longer term (~ ten years or more), where the former case corresponds to Lunar or Mars cargo-type transfer vehicles and the latter case corresponds to outer solar system explorations. EP based systems are hence important to make efficient interplanetary transfer in the inner solar system as well as to realize long-distant cruise to reach outer planets. For these purposes, technology expansion is really required not only for high-power and large-scale applications but also for smaller and low power challenges.

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