Special Issue "Electric Propulsion"

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

Deadline for manuscript submissions: closed (30 September 2020).

Special Issue Editor

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

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Published Papers (15 papers)

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Research

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Article
Background Pressure Effects on the Performance of a 20 kW Magnetically Shielded Hall Thruster Operating in Various Configurations
Aerospace 2021, 8(3), 69; https://doi.org/10.3390/aerospace8030069 - 09 Mar 2021
Viewed by 574
Abstract
The paper reports the characterization results of a 20 kW-class magnetically shielded Hall thruster in three different configurations and operating with a centrally mounted cathode. The characterization was carried out at two different pumping speeds in SITAEL’s IV10 vacuum chamber, resulting in two [...] Read more.
The paper reports the characterization results of a 20 kW-class magnetically shielded Hall thruster in three different configurations and operating with a centrally mounted cathode. The characterization was carried out at two different pumping speeds in SITAEL’s IV10 vacuum chamber, resulting in two different background pressure levels for each tested operating point. A linear behavior of discharge current and thrust values versus the anode mass flow rate was noticed for both pumping speeds levels and for all the three configurations. In addition, the thrust and discharge current values were always found to be lower at lower background pressure levels. From the performance levels, a preliminary estimate of the ingested mass flow rates was performed, and the values were then compared to a recently developed background flow model. The results suggested that, for this thruster and in the tested operating regimes, the change in performance due to background pressure could be ascribed not only to the ingestion of external mass flow coming from the chamber but also to other physical processes caused by the flux of residual background neutrals. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Polyatomic Ion-Induced Electron Emission (IIEE) in Electrospray Thrusters
Aerospace 2020, 7(11), 153; https://doi.org/10.3390/aerospace7110153 - 24 Oct 2020
Cited by 3 | Viewed by 1335
Abstract
To better characterize the lifetime and performance of electrospray thrusters, electron emission due to electrode impingement by the propellant cation 1-ethyl-3-methylimidazolium (EMI+) has been evaluated with semi-empirical modeling techniques. Results demonstrate that electron emission due to grid impingement by EMI+ [...] Read more.
To better characterize the lifetime and performance of electrospray thrusters, electron emission due to electrode impingement by the propellant cation 1-ethyl-3-methylimidazolium (EMI+) has been evaluated with semi-empirical modeling techniques. Results demonstrate that electron emission due to grid impingement by EMI+ cations becomes significant once EMI+ attains a threshold velocity of ∼9×105 cm s1. The mean secondary electron yield, γ¯, exhibits strong linearity with respect to EMI+ velocity for typical electrospray operating regimes, and we present a simple linear fit equation corresponding to thruster potentials greater than 1 kV. The model chosen for our analysis was shown to be the most appropriate for molecular ion bombardments and is a useful tool in estimating IIEE yields in electrospray devices for molecular ion masses less than ∼1000 u and velocities greater than ∼106 cm s1. Droplet-induced electron emission (DIEE) in electrospray thrusters was considered by treating a droplet as a macro-ion, with low charge-to-mass ratio, impacting a solid surface. This approach appears to oversimplify back-spray phenomena, meaning a more complex analysis is required. While semi-empirical models of IIEE, and the decades of solid state theory they are based upon, represent an invaluable advance in understanding secondary electron emission in electrospray devices, further progress would be gained by investigating the complex surfaces the electrodes acquire over their lifetimes and considering other possible emission processes. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Emission Modes in Electrospray Thrusters Operating with High Conductivity Ionic Liquids
Aerospace 2020, 7(10), 141; https://doi.org/10.3390/aerospace7100141 - 25 Sep 2020
Cited by 2 | Viewed by 1202
Abstract
Electrospray thruster life and mission performance are strongly influenced by grid impingement, the extent of which can be correlated with emission modes that occur at steady-state extraction voltages, and thruster command transients. Most notably, we experimentally observed skewed cone-jet emission during steady-state electrospray [...] Read more.
Electrospray thruster life and mission performance are strongly influenced by grid impingement, the extent of which can be correlated with emission modes that occur at steady-state extraction voltages, and thruster command transients. Most notably, we experimentally observed skewed cone-jet emission during steady-state electrospray thruster operation, which leads to the definition of an additional grid impingement mechanism that we termed “tilted emission”. Long distance microscopy was used in conjunction with high speed videography to observe the emission site of an electrospray thruster operating with an ionic liquid propellant (EMI-Im). During steady-state thruster operation, no unsteady electrohydrodynamic emission modes were observed, though the conical meniscus exhibited steady off-axis tilt of up to 15°. Cone tilt angle was independent over a wide range of flow rates but proved strongly dependent on extraction voltage. For the geometry and propellant used, the optimal extraction voltage was near 1.6 kV. A second experiment characterized transient emission behavior by observing startup and shutdown of the thruster via flow or voltage. Three of the four possible startup and shutdown procedures transition to quiescence within ∼475 μs, with no observed unsteady modes. However, during voltage-induced thruster startup, unsteady electrohydrodynamic modes were observed. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Future Directions for Electric Propulsion Research
Aerospace 2020, 7(9), 120; https://doi.org/10.3390/aerospace7090120 - 20 Aug 2020
Cited by 9 | Viewed by 2055
Abstract
The research challenges for electric propulsion technologies are examined in the context of s-curve development cycles. It is shown that the need for research is driven both by the application as well as relative maturity of the technology. For flight qualified systems such [...] Read more.
The research challenges for electric propulsion technologies are examined in the context of s-curve development cycles. It is shown that the need for research is driven both by the application as well as relative maturity of the technology. For flight qualified systems such as moderately-powered Hall thrusters and gridded ion thrusters, there are open questions related to testing fidelity and predictive modeling. For less developed technologies like large-scale electrospray arrays and pulsed inductive thrusters, the challenges include scalability and realizing theoretical performance. Strategies are discussed to address the challenges of both mature and developed technologies. With the aid of targeted numerical and experimental facility effects studies, the application of data-driven analyses, and the development of advanced power systems, many of these hurdles can be overcome in the near future. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Lifetime Considerations for Electrospray Thrusters
Aerospace 2020, 7(8), 108; https://doi.org/10.3390/aerospace7080108 - 29 Jul 2020
Cited by 4 | Viewed by 1807
Abstract
Ionic liquid electrospray thrusters are capable of producing microNewton precision thrust at a high thrust–power ratio but have yet to demonstrate lifetimes that are suitable for most missions. Accumulation of propellant on the extractor and accelerator grids is thought to be the most [...] Read more.
Ionic liquid electrospray thrusters are capable of producing microNewton precision thrust at a high thrust–power ratio but have yet to demonstrate lifetimes that are suitable for most missions. Accumulation of propellant on the extractor and accelerator grids is thought to be the most significant life-limiting mechanism. In this study, we developed a life model to examine the effects of design features, operating conditions, and emission properties on the porous accelerator grid saturation time of a thruster operating in droplet emission mode. Characterizing a range of geometries and operating conditions revealed that modifying grid aperture radius and grid spacing by 3–7% can significantly improve thruster lifetime by 200–400%, though a need for explicit mass flux measurement was highlighted. Tolerance analysis showed that misalignment can result in 20–50% lifetime reduction. In addition, examining the impact of electron backstreaming showed that increasing aperture radius produces a significant increase in backstreaming current compared to changing grid spacing. A study of accelerator grid bias voltages revealed that applying a reasonably strong accelerator grid potential (in the order of a kV) can minimize backstreaming current to negligible levels for a range of geometries. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
UWE-4: First Electric Propulsion on a 1U CubeSat—In-Orbit Experiments and Characterization
Aerospace 2020, 7(7), 98; https://doi.org/10.3390/aerospace7070098 - 17 Jul 2020
Cited by 4 | Viewed by 1657
Abstract
The electric propulsion system NanoFEEP was integrated and tested in orbit on the UWE-4 satellite, which marks the first successful demonstration of an electric propulsion system on board a 1U CubeSat. In-orbit characterization measurements of the heating process of the propellant and the [...] Read more.
The electric propulsion system NanoFEEP was integrated and tested in orbit on the UWE-4 satellite, which marks the first successful demonstration of an electric propulsion system on board a 1U CubeSat. In-orbit characterization measurements of the heating process of the propellant and the power consumption of the propulsion system at different thrust levels are presented. Furthermore, an analysis of the thrust vector direction based on its effect on the attitude of the spacecraft is described. The employed heater liquefies the propellant for a duration of 30 min per orbit and consumes 103 ± 4 mW. During this time, the respective thruster can be activated. The propulsion system including one thruster head, its corresponding heater, the neutralizer and the digital components of the power processing unit consume 8.5 ± 0.1 mW · μ A−1 + 184 ± 8.5 mW and scales with the emitter current. The estimated thrust directions of two thruster heads are at angles of 15.7 ± 7.6 and 13.2 ± 5.5 relative to their mounting direction in the CubeSat structure. In light of the very limited power on a 1U CubeSat, the NanoFEEP propulsion system renders a very viable option. The heater of subsequent NanoFEEP thrusters was already improved, such that the system can be activated during the whole orbit period. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Electrospray Propulsion Engineering Toolkit (ESPET)
Aerospace 2020, 7(7), 91; https://doi.org/10.3390/aerospace7070091 - 04 Jul 2020
Viewed by 1666
Abstract
We report on the development of a software tool, the Electrospray Propulsion Engineering Toolkit (ESPET), that is currently being shared as a web application with the purpose to accelerate the development of electrospray thruster arrays for space propulsion. ESPET can be regarded as [...] Read more.
We report on the development of a software tool, the Electrospray Propulsion Engineering Toolkit (ESPET), that is currently being shared as a web application with the purpose to accelerate the development of electrospray thruster arrays for space propulsion. ESPET can be regarded as a database of microfluidic properties and electrohydrodynamic scaling models that are combined into a performance estimation tool. The multiscale model integrates experimental high-level physics characterization of microfluidic components in a full-scale electrospray propulsion (ESP) microfluidic network performance solution. ESPET takes an engineering model approach that breaks the ESP system down into multiple microfluidic components or domains that can be described by either analytical microfluidic or reduced order numerical solutions. ESPET can be divided into three parts: a central database of critical microfluidic properties, a microfluidic domain modeler, and a microfluidic network solver. Two options exist for the network solution, a detailed multi-domain solver and a QuickSolver designed for rapid design and testing of simple three-domain reservoir-feed-emitter arrays. The multi-domain network solver exploits the Hagen–Poiseuille/Ohm’s law analogy by using the publicly available SPICE (Simulation Program with Integrated Circuit Emphasis) electric circuit simulation software to solve the flow properties of the microfluidic network. Both the multi-domain and QuickSolver solutions offer Monte Carlo analysis of arrays based on user supplied tolerances on design parameters. Benchmarking demonstration examples are provided for experimental work in the literature, as well as recent experimental work conducted at Busek Co. The demonstration examples include ionic liquid propelled systems using active and passive capillary emitters, externally wetted emitter needles, and porous glass emitters, as well as a liquid metal system based on an externally wetted emitter needle. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Article
Characterization of Vacuum Arc Thruster Performance in Weak Magnetic Nozzle
Aerospace 2020, 7(6), 82; https://doi.org/10.3390/aerospace7060082 - 19 Jun 2020
Cited by 3 | Viewed by 1648
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. 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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Article
Sequential Low-Thrust Orbit-Raising of All-Electric Satellites
Aerospace 2020, 7(6), 74; https://doi.org/10.3390/aerospace7060074 - 04 Jun 2020
Viewed by 1901
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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Article
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
Viewed by 1566
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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Article
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
Cited by 1 | Viewed by 1991
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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Article
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
Cited by 1 | Viewed by 2424
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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Review
Electric Propulsion Methods for Small Satellites: A Review
Aerospace 2021, 8(1), 22; https://doi.org/10.3390/aerospace8010022 - 18 Jan 2021
Cited by 1 | Viewed by 1717
Abstract
Over 2500 active satellites are in orbit as of October 2020, with an increase of ~1000 smallsats in the past two years. Since 2012, over 1700 smallsats have been launched into orbit. It is projected that by 2025, there will be 1000 smallsats [...] Read more.
Over 2500 active satellites are in orbit as of October 2020, with an increase of ~1000 smallsats in the past two years. Since 2012, over 1700 smallsats have been launched into orbit. It is projected that by 2025, there will be 1000 smallsats launched per year. Currently, these satellites do not have sufficient delta v capabilities for missions beyond Earth orbit. They are confined to their pre-selected orbit and in most cases, they cannot avoid collisions. Propulsion systems on smallsats provide orbital manoeuvring, station keeping, collision avoidance and safer de-orbit strategies. In return, this enables longer duration, higher functionality missions beyond Earth orbit. This article has reviewed electrostatic, electrothermal and electromagnetic propulsion methods based on state of the art research and the current knowledge base. Performance metrics by which these space propulsion systems can be evaluated are presented. The article outlines some of the existing limitations and shortcomings of current electric propulsion thruster systems and technologies. Moreover, the discussion contributes to the discourse by identifying potential research avenues to improve and advance electric propulsion systems for smallsats. The article has placed emphasis on space propulsion systems that are electric and enable interplanetary missions, while alternative approaches to propulsion have also received attention in the text, including light sails and nuclear electric propulsion amongst others. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Review
State-of-the-Art and Advancement Paths for Inductive Pulsed Plasma Thrusters
Aerospace 2020, 7(8), 105; https://doi.org/10.3390/aerospace7080105 - 24 Jul 2020
Cited by 3 | Viewed by 2217
Abstract
An inductive pulsed plasma thruster (IPPT) operates by pulsing high current through an inductor, typically a coil of some type, producing an electromagnetic field that drives current in a plasma, accelerating it to high speed. The IPPT is electrodeless, with no direct electrical [...] Read more.
An inductive pulsed plasma thruster (IPPT) operates by pulsing high current through an inductor, typically a coil of some type, producing an electromagnetic field that drives current in a plasma, accelerating it to high speed. The IPPT is electrodeless, with no direct electrical connection between the externally applied pulsed high-current circuit and the current conducted in the plasma. Several different configurations were proposed and tested, including those that produce a plasma consisting of an accelerating current sheet and those that use closed magnetic flux lines to help confine the plasma during acceleration. Specific impulses up to 7000 s and thrust efficiencies over 50% have been measured. The present state-of-the-art for IPPTs is reviewed, focusing on the operation, modeling techniques, and major subsystems found in various configurations. Following that review is documentation of IPPT technology advancement paths that were proposed or considered. Full article
(This article belongs to the Special Issue Electric Propulsion)
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Review
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
Cited by 4 | Viewed by 2313
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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