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Many plasma systems and technologies, such as Hall thrusters for spacecraft propulsion, exhibit complex underlying physics that affect the global operation. When characterizing such systems in an experiment, obtaining full spatiotemporal maps of the involved state variables can be, thus, highly informative. However, this goal is not practically realizable because of various experimental limitations, e.g., finite spatial resolution of the diagnostics and geometrical accessibility constraints. Therefore, having the capability to reconstruct the full high-dimensional states of plasma systems from low-dimensional time-history measurements is greatly desirable. Compressed sensing is a signal processing technique that can answer this crucial need. However, existing compressed sensing approaches have several limitations that restrict their effectiveness for complex physical systems like plasma technologies. These include the need for abundant sensor measurements and a principled sensor placement. In this paper, we demonstrate the capabilities of Shallow Recurrent Decoder (SHRED) architecture for compressed sensing. We show in several plasma test cases that SHRED can robustly infer full high-dimensional spatiotemporal state vectors of these systems (i.e., all macroscopic plasma properties) from minimal system information. This minimal information can consist of three finite time-history measurements of either local values of a plasma property or the global plasma properties (spatially averaged or performance parameters). An application of SHRED's inference capability in the numerical plasma simulation context is "super-resolution" enhancement. We will discuss this application by presenting how SHRED can effectively establish mappings between a low-resolution and a high-resolution simulation, recovering detailed spatial plasma features that are below the simulation's grid size. [ABSTRACT FROM AUTHOR]

Ionic liquid electrospray thrusters represent an alternative propulsion method for spacecraft to conventional plasma propulsion because they do not require plasma generation, which significantly increases the thrust efficiency. The porous emitter thruster has the advantages of simple propellant feeding and multi-site emissions, which miniaturize the thruster size and increase thrust. However, the multi-scale nature, that is, nano- to micrometer-sized menisci on the millimeter-size porous needle tip, makes modeling multi-site emissions difficult, and direct observation is also challenging. This paper proposes a simple model for multi-site emissions, which assumes that the ionic conductivity or ion transport in the porous media determines the ion-emission current. The conductivity was evaluated by comparing the experimental and numerical data based on the model. The results suggest that the ionic conductivity of the porous emitter is suppressed by the ion–pore wall friction stress. Additionally, the model indicates that the emission area expansion on the porous emitter creates the unique curve shape of the current vs voltage characteristics for multi-site emissions. [ABSTRACT FROM AUTHOR]

A multi-scale simulation framework for ion–solid interactions in plasma-exposed materials provides crucial insight into advancing fusion energy and space electric propulsion. Leveraging binary-collision approximation (BCA) simulations, the framework uniquely predicts sputter yields and analyzes material transport within volumetrically complex materials. This approach, grounded in the validated BCA code TRI3DYN, addresses key limitations in existing models by accurately capturing ion–solid interaction physics. A case study is presented, highlighting the framework's ability to replicate experimental sputter yield results, underscoring its reliability and potential for designing durable materials in harsh plasma environments. Insights into sputtering transport phenomenology mark a significant advancement in material optimization for improved resilience in plasma-facing applications. [ABSTRACT FROM AUTHOR]

Field-reversed configuration electric propulsion is an advanced space electromagnetic propulsion technology with significant application prospects but poor thrust performance. To delve into the intrinsic physical mechanisms, a model of the rotating magnetic field penetrating into the plasma and azimuthal electron current driving was developed. The simulation results show that rotating magnetic field strength, frequency, and initial seed plasma density are the key factors that exist as an optimal threshold. Specifically, the rotating magnetic field feed current (i.e., magnetic field strength) was not less than 1000 A, the rotating magnetic field frequency was ∼200–300 kHz, and the plasma density was approximately 1 × 1018 m−3 order of magnitude. [ABSTRACT FROM AUTHOR]

Hall thrusters are one of the most successful and prevalent electric propulsion systems for spacecraft in use today. However, they are also complex devices and their unique E × B configuration makes modeling of the underlying plasma discharge challenging. In this work, a steady-state model of a Hall thruster is developed and a complete analytical solution presented that is shown to be in reasonable agreement with experimental measurements. A characterization of the discharge shows that the peak plasma density and ionization rate nearly coincide and both occur upstream of the peak electric field. The peak locations also shift as the thruster operating conditions are varied. Three key similarity parameters emerge that govern the plasma discharge and which are connected via a thruster current–voltage relation: a normalized discharge current, a normalized discharge voltage, and an amalgamated parameter, α ¯ , that contains all system geometric and magnetic field information. For a given normalized discharge voltage, the similarity parameter α ¯ must lie within a certain range to enable high thruster performance. When applied to a krypton thruster, the model shows that both the propellant mass flow rate and the magnetic field strength must be simultaneously adjusted to achieve similar efficiency to a xenon thruster (for the same thruster geometry, discharge voltage, and power level). [ABSTRACT FROM AUTHOR]

Laser Thomson scattering (LTS) is a measurement technique that can determine electron velocity distribution functions in plasma systems. However, accurately inferring quantities of interest from an LTS signal requires the selection of a plasma physics submodel, and comprehensive uncertainty quantification (UQ) is needed to interpret the results. Automated model selection, parameter estimation, and UQ are particularly challenging for low-density, low-temperature, potentially non-Maxwellian plasmas like those created in space electric propulsion devices. This paper applies Bayesian inference and model selection to a Raman-calibrated LTS diagnostic in the context of such plasmas. Synthetic data are used to explore the performance of the method across signal-to-noise ratios and model fidelity regimes. Plasmas with Maxwellian and non-Maxwellian velocity distributions are well characterized using priors that span a range of accuracy and specificity. The model selection framework is shown to accurately detect the type of plasmas generating the electron velocity distribution submodel for signal-to-noise ratios greater than around 5. In addition, the Bayesian framework validates the widespread use of 95% confidence intervals from least-squares inversion as a conservative estimate of the uncertainty bounds. However, epistemic posterior correlations between the variables diverge between least-squares and Bayesian estimates as the number of variable parameters increases. This divergence demonstrates the need for Bayesian inference in cases where accurate correlations between electron parameters are necessary. Bayesian model selection is then applied to experimental Thomson scattering data collected in a nanosecond pulsed plasma, generated with a discharge voltage of 5 and 10 kV at a neutral argon background pressure of 7 Torr-Ar. The Bayesian maximum a posteriori estimates of the electron temperature and number density are 1.98 and 2.38 eV and 2.6 × 1018 and 2.72 × 1018 m−3, using the Maxwellian and Druyvesteyn submodels, respectively. Furthermore, for this dataset, the model selection criterion indicates strong support for the Maxwellian distribution at 10 kV discharge voltage and no strong preference between Maxwellian and Druyvesteyn distributions at 5 kV. The logarithmic Bayes' factors for these cases are −35.76 and 1.07, respectively. [ABSTRACT FROM AUTHOR]

The Alternative Low-Power Hybrid Ion Engine (alphie) is a new technology for space propulsion based on plasma. Its distinct characteristic is the counterflow of charges (ions and electrons) passing through its two-grid system. This means that electrons coming from an external cathode are accelerated toward the ionization chamber, in which a neutral gas (typically Ar) is injected. The strong magnetic field therein confines these electrons, which ionize and exchange energy with the propellant gas. Thus, the operation of alphie is strongly affected by the electrons coming from the external cathode and their collisions with the neutral atoms. This work studies the counterflow employing a particle-in-cell simulation of ions and electrons passing through a single hole as a function of the electron cathode currents (Ice) and potential drops between grids (VAC). Transparency of the grid system to ions and electrons and the ion current extracted by the grid system are studied under sweeps of these two parameters. The number of ionization events by each high-energy electron entering the ionization chamber is evaluated using a physical model based on the gas density and the cross section for ionization. These new ions are then extracted by the same electric field that accelerates the electrons inward. Thus, simulations are self-consistent, since the ionizing electron flow from the external cathode drives the ion outflow at the exit section of the two-grid system. The electrical transparency of the two-grid system to ions and electrons, related to the axial charge currents, is also studied under sweeps of aforementioned operation parameters. This new way to deal with ionizations can be useful to study other plasma thrusters in which electrons for ionization come from an external cathode without modeling the complex structure of the ionization chamber. [ABSTRACT FROM AUTHOR]

Ionic-liquid ion sources produce beams of charged particles through evaporation and acceleration of ions and charged droplets from the surface of an ionic liquid. The composition of the emitted beam can impact the performance of ion sources for various applications such as focused beams for microfabrication and space propulsion. Numerical inference is considered for quantification of the beam composition of an ionic-liquid ion source through determining the current fraction of different species along with providing uncertainty in inferred values. An analysis of previously presented data demonstrates the ability to quantify the presence of ion clusters, including the distinct presence of heavy ion clusters such as heptamers. Quantification of beam composition will be an important technique for quantitative comparison of different time-of-flight data. [ABSTRACT FROM AUTHOR]

The quad confinement plasma source is a novel plasma device developed for space propulsion applications, whose core is an E × B discharge with open electron drift. The magnetic field is produced by independently powered electromagnets able to generate different magnetic field topologies with the ultimate aim of manipulating the ion flow field for achieving thrust vectoring. In this work, we map the ion velocity in the plasma ejected from the quad confinement thruster with different magnetic configurations using non-intrusive laser-induced fluorescence diagnostics. Measurements show a steep ion acceleration layer located 8 cm downstream the exit plane of the discharge channel, detached from any physical boundary of the plasma source. In this location, the ion velocity increases from 3 to 10 km/s within a 1 cm axial region. The ion acceleration profile has been characterized under multiple testing conditions in order to identify the influence of the magnetic field intensity and topology on this peculiar ion acceleration layer. [ABSTRACT FROM AUTHOR]

Cathodic arcs (CA) are important across a wide range of applications including thin film deposition and solid fuel space propulsion. Understanding how cathode spots (CSs) move in magnetic fields has important implications for the deposition rates and thrust delivered as well as cathode/fuel usage efficiency. In this work, we investigated the dynamics of cathode spot (CS) motion in the presence of various axial magnetic fields (AMFs) in a high-current multi-spot CA system. Two photography techniques were used (i) long exposure and (ii) high-speed framing photography. The long-exposure images of cathode spots suggested that the cathode spots do not travel in spiral paths produced by Robson drift, as observed in previous vacuum interrupter system experiments. The reduced prevalence of Robson drift was found to be insufficiently explained considering only anodic plasma emission difference arising from the annular anode geometrical setup. High-speed framing images of cathode spots during discharges revealed a statistically significant reduction of the speed of cathode spot motion, as characterized by a decrease in K value as the applied AMF increases. [ABSTRACT FROM AUTHOR]

Direct numerical simulations (DNS) are conducted to analyze the effect of recess in the mixing and combustion performance of gaseous methane—oxygen injection systems. The recess length is varied from 0 to 2 injector diameters to analyze the sensitivity of various physical processes to this geometrical feature. It is found that the injection recess enhances the development of Kelvin–Helmholtz instabilities (KHI), due to the more consistent jet flow disposition at the injection plane, and the higher velocities. A Strouhal number Str R ≈ 0.22 , based on the injector diameter and on the average reacting jet velocity is found for all simulations. In addition, it is found that the enhancement of KHI acts as an injection of turbulent kinetic energy in the large scales, improving macroscale mixing and combustion performance. Finally, the normalized enhancements in combustion performance and averaged hydroxyl emissivity are found to be in good agreement with experimental results in a similar setup. [ABSTRACT FROM AUTHOR]

In this work, we investigate the properties of the plasma induced by focusing a high-power laser beam on an aluminum target that was cooled by a helium refrigerator from room temperature down to 20 K. Fast, streak photographs of the plasma were taken at different temperatures and laser energies. From the images obtained, position-vs-time plots were made for each experiment, and from them, the speed was calculated. Additionally, narrowband interference filters were employed to image the dynamics of ions and neutrals separately. It was found that the plasma plume has two distinct speeds: that of its center and that of the outer edge. For unfiltered images, the former has values within the interval 6.2 to 9.1 km/s, while the latter can reach speeds of the order of hundreds of km/s. It was found that the plume of a target cooled to 20 K has a length that is 8%–12% less than the corresponding size at room temperature. Chilling the target did not seem to affect significantly either the plume's speed of expansion or the size of the crater produced. Lower bounds were estimated for the momentum imparted to the ejecta and the specific impulse. The latter can reach a 920-s level, nearly twice as much the amount obtained with chemical rocket fuel. [ABSTRACT FROM AUTHOR]

Applied field magnetoplasmadynamic thrusters (AF-MPDTs) are one of the potential thrusters for high power space propulsion. However, the high anode power deposition limits the improvement in the thruster's efficiency. According to previous research studies, the anode voltage drop is supposed to be the main factor of the anode power deposition in AF-MPDTs without considering the effect of the centrifugal force. Actually, the centrifugal force cannot be neglected when the magnetic field is high, where the anode voltage drop is not the only dominant factor of the anode power deposition. Due to the significant centrifugal force, there might be no obvious shortage of charge carriers near the anode surface in AF-MPDTs with propellant injection between the cathode and anode. The electrons will be compressed to the anode surface instead of the cathode. In order to investigate the reason for the anode power deposition in AF-MPDTs, a model considering the effect of the centrifugal force is established. The calculated anode power deposition shows good agreement with the corresponding experimental data. In addition, when the magnetic field is high, the viscous heating cannot be neglected, which is as significant as the anode sheath heating and Joule heating. Finally, in order to reduce the anode power deposition, a magnetic field distribution with a high value near the cathode and a low value near the anode is proposed. According to the evaluation by the model, the anode power deposition can be reduced by 30 % via the new magnetic field when the magnetic field exceeds 0.1 T. [ABSTRACT FROM AUTHOR]

Turbulence in hollow cathodes used for space propulsion is believed to play an important role in anomalous electron transport and ion heating. In this work, the implementation of coherent Thomson scattering to identify and characterize MHz-frequency ion acoustic turbulence and kHz-frequency oscillations in the plume of a hollow cathode is achieved. In the presence of a background magnetic field of a Hall thruster, a number of unstable modes are observed. A directive ion acoustic mode propagating predominantly within a restricted angle around the magnetic field is found, exhibiting an energy scaling with wavenumber k of the form k − 5.2 ± 0.58 , which differs from the classic Kadomtsev k − 3 scaling for unmagnetized conditions. Bi-directional ion acoustic mode fluctuations propagating over a range of angles with respect to the magnetic field have been measured, possibly signifying the existence of a large-amplitude plasma wave, similar to the Buneman instability. Finally, electron density fluctuations in the kHz-frequency range, a possible consequence of drift-driven instabilities in the plane perpendicular to the magnetic field, have also been identified. These results not only are an indication of the diversity of wave types that exist in hollow cathode plumes but also point to the key role played by the presence of, and the configuration of, the magnetic field in their appearance. [ABSTRACT FROM AUTHOR]

Ionic liquid ion sources are expected to be used in a wide range of applications such as space electric propulsion and focused ion beam micromachining. It is known that the backstreaming of secondary charged species generated by ion beam impacts can cause unexpected temperature rise and chemical changes in ionic liquids. This paper reports on results of heating experiments using a sharp needle emitter wetted with an ionic liquid, 1-ethyl-3-methyl imidazolium bis(trifluoromethanesulfonyl)amide, at temperatures in a range from room temperature to 120 °C. Current measurements show that positive and negative electrospray currents from the heated emitter increased as the temperature increased. Time-of-flight (TOF) mass spectrometric measurements reveal that the beam composition changed significantly with increasing temperature, indicating that charged droplets as well as ions were emitted from the heated emitter. The TOF data show that a significant fraction of the current is due to droplets at higher temperatures. On the basis of the results obtained, the size and charge of the emitted droplets are discussed. The beam is roughly estimated to contain charged droplets with a diameter of around 20 nm at 120 °C. [ABSTRACT FROM AUTHOR]

Ion-induced secondary electron emission (SEE) is a fundamental surface interaction that strongly influences many plasma discharges. Recently, interest in iodine plasmas is growing due to new material processing and space propulsion applications, but data for SEE yields due to iodine ion bombardment remain scarce. Additionally, due to the formation of multiple ion species in typical iodine plasmas and surface chemical reactions leading to iodide layer formation, the effective SEE yield is expected to differ from that for individual ion species on clean surfaces. In this work, we measure the SEE yield of multi-species iodine ion beams bombarding different target materials (Mo, W, Al, Ti, Cu, carbon-carbon, and steel) in the energy range of 0.6–1.4 keV. An ion beam is produced by extracting and accelerating ions from a gridded ion source based on an inductively coupled plasma (ICP). SEE yields of downstream targets are measured using a conventional electrostatic probe technique, and the ion beam composition is determined using time-of-flight spectrometry. The beam is composed predominately of atomic ( I + ) and molecular ( I 2 + ) ions whose ratio changes depending on the ICP power. Yields depend strongly on the target material and beam composition and vary between 0.05 and 0.4 depending on whether potential or kinetic emission processes dominate. [ABSTRACT FROM AUTHOR]

Low-power (from the sub-kilowatt range up to a few kilowatts) inductively coupled plasma (ICP) electrothermal thrusters for space propulsion are potential alternatives to low-power arcjet thrusters, which are often implemented on geostationary satellites for north–south station keeping. We develop an axisymmetric two-dimensional magnetohydrodynamic (MHD) numerical simulation technique using a two-temperature plasma model as a computer-aided engineering tool for low-power argon ICP electrothermal thrusters. Furthermore, to the best of our knowledge, this is the first study to provide a basic understanding of the plasma-fluid characteristics of low-power ICP electrothermal thrusters. To this end, we perform the MHD numerical simulation for a low-power argon ICP electrothermal thruster model that was developed for thrust measurement experiments to validate the proposed numerical simulation technique. The numerical results indicate that the plasma flow produced in the low-power argon ICP electrothermal thruster model is basically in a strongly thermal and ionizing nonequililbrium state. In addition, the numerical results show that the experimentally measured thrust forces can be reproduced with an underestimation of 10% or less over the RF input power range of 0–0.6 kW considered in the experiment. Moreover, the numerical results suggest that the main reason why the performance of the low-power argon ICP electrothermal thruster model is significantly lower than the target performance is the considerable heat loss of more than 90% of the input power through the flow-channel wall. [ABSTRACT FROM AUTHOR]

The feasibility of helicon plasma thrusters, commonly considered the most promising for deep space propulsion, is evaluated in terms of their plasma production efficiency, longevity, operation stability, and need for massive magnets with large power consumption. It is shown that helicon thrusters should be inferior to those based on inductively coupled plasmas. [ABSTRACT FROM AUTHOR]

Turbulent mixing in a methane–oxygen recessed injector is studied using direct numerical simulations. The operating point is chosen to be fuel-rich and at high pressure to recreate a representative environment for space propulsion applications. The results are used to investigate the transport of the turbulent mixture fraction statistics and the validity of conventional transport models. It is observed that molecular diffusion is only relevant near the boundary layer of the injection recess cavity and at the recirculation zone. Moreover, turbulent mixing in the axial direction is negligible as radial turbulent diffusion dominates. Radial turbulent diffusion near injection is driven by Kelvin–Helmholtz Instabilities (KHI) manifesting at large scales in the order of the injector geometry. The dominance of this process over microscale mixing originates negative turbulent diffusion, which produces a mixture resegregation and the appearance of lean pockets far from the oxidizer injection plane. Gradient models display poor capabilities for the prediction of this sort of phenomena. Closure models for the turbulent mixing transport terms are proposed and evaluated. An anisotropic gradient model is devised, providing performance improvements within the recess cavity and the recirculation region. In addition, a novel filtered Reynolds-averaged Navier Stokes approach based on the mixing state is proposed. This new methodology shows excellent prediction capabilities in the regions dominated by KHI, accurately predicting negative turbulent diffusivity. The challenges associated with this model are commented on, and strategies to enable its application are proposed. [ABSTRACT FROM AUTHOR]

Low-temperature E × B plasmas are used in various applications, such as Hall thrusters for satellite propulsion, ion sources and magnetron discharges for plasma processing, and negative ion sources for neutral beam injection in fusion. The plasmas in these devices are partially magnetized, meaning that the electrons are strongly magnetized while the ions are not. They are subject to various micro- and macro-instabilities that differ significantly from instabilities in fusion plasmas. These instabilities are often triggered by the large difference in electron and ion drift velocities in the E × B direction. The possibility of maintaining a large electric field in the quasineutral plasma of Hall thrusters despite anomalous electron transport, or the presence of strong double layers associated with the azimuthal rotation of plasma structures ("rotating spokes") in magnetron discharges and Hall thrusters are examples of the very challenging and exciting physics of E × B devices. The turbulence and instabilities present in E × B plasma devices constitute a major obstacle to the quantitative description of these devices and to the development of predictive codes and are the subject of intense research efforts. In this tutorial, we discuss the key aspects of the physics of low-temperature partially magnetized E × B plasmas, as well as recent advances made through simulations, theory, and experiments in our understanding of the various types of instabilities (such as gradient-drift/Simon-Hoh and lower hybrid instabilities, rotating ionization waves, electron cyclotron drift instability, modified two-stream instability, etc.) that occur in these plasmas. [ABSTRACT FROM AUTHOR]

Space propulsion is unique among many proposed applications of the dense plasma focus in being critically dependent on the availability of a scaling theory that is well-grounded in physics, in conformity with existing experimental knowledge and applicable to experimentally untested configurations. This paper derives such a first-principles-based scaling theory and illustrates its application to a novel space propulsion concept, where the plasma focus sheath is employed as a power density amplifying mechanism to transport electric energy from a capacitive storage to a current-driven fusion load. For this purpose, a Generalized Plasma Focus problem is introduced and formulated. It concerns a finite, axisymmetric plasma, driven through a neutral gas at supersonic speed over distances much larger than its typical gradient scale length by its azimuthal magnetic field while remaining connected with its pulse power source through suitable boundaries. The Gratton-Vargas equation is rederived from the scaling properties of the equations governing plasma dynamics and solved for algebraically defined initial (insulator) and boundary (anode) surfaces. Scaling relations for a new space propulsion concept are derived. This consists of a modified plasma focus with a tapered anode that transports current from a pulsed power source to a consumable portion of the anode in the form of a hypodermic needle tube continuously extruded along the axis of the device. When the tube is filled with deuterium, the device may serve as a small-scale version of magnetized liner inertial fusion (MAGLIF) that could avoid failure of neutron yield scaling in a conventional plasma focus. [ABSTRACT FROM AUTHOR]

Pendulum thrust stands are used to measure the thrust of electric propulsion systems for spacecraft. A thruster is mounted on a pendulum and operated, and the pendulum displacement due to thrust is measured. In this type of measurement, the pendulum is also affected by nonlinear tensions due to wiring and piping that deteriorate the accuracy of the measurement. This influence cannot be ignored in high power electric propulsion systems because complicated piping and thick wirings are required. Therefore, to reduce the influence of tension due to wires and tubes, we developed an inverted pendulum-type thrust stand with pipes and wirings as springs. In this paper, we first derive the design guidelines for spring-shaped wires; the necessary conditions for sensitivity, responsivity, spring shape, and electric wire were formulated. Next, a thrust stand was designed and fabricated based on these guidelines, and the performance of the stand was evaluated through calibration and thrust measurements using a 1 kW-class magneto-plasma-dynamics thruster. The sensitivity of the thrust stand was 17 mN/V, the normalized standard deviation of the variation of the measured values owing to the structure of the thrust stand was 1.8 × 10−3, and the thermal drift during the long-time operation was ∼4.5 × 10−3 mN/s. [ABSTRACT FROM AUTHOR]

The usage of direct numerical simulations for research of turbulent combustion for space propulsion applications is explored. With this goal in mind, the combustion near the injection of a fuel-rich methane–oxygen flame at 20 bar is simulated using a massively parallelized solver. The statistical properties of the relevant physical fields are examined to study interactions between turbulence and combustion. This analysis is complemented by an investigation and quantification of the error sources in direct numerical simulations of turbulent diffusion flame. A method to estimate the statistical error is derived based on the classical inference theory. In addition, critical resolution criteria are discussed using a mesh sensitivity analysis. [ABSTRACT FROM AUTHOR]

A torsional thrust balance has been designed and validated by Surrey Space Centre and Added Value Solutions UK Ltd. in collaboration with the UK Space Agency. The thrust stand has been tested with two electric propulsion (EP) systems operating with xenon: the Halo thruster and the XJET thruster. The first consists of a low-power (<1 kW) Hall effect-based thruster, whose thrust level is between 3 and 20 mN, depending on the power of the system. The second is an electron cyclotron resonance thruster whose operative point is in the 0.3–1.5 mN thrust range. The thruster is mounted on a titanium rotating beam, whose movement is measured by an optical fiber displacement sensor. The thrusters' direct current electrical connections are routed through room temperature liquid metal pots and microwave power is transmitted via a wireless transfer system, minimizing friction effects. To reduce thermal issues during long thruster operations, the torsional thrust balance is designed with a water-cooling hub around the flex pivot. Noise from the laboratory environment is lessened by using four vibration-dampening spring systems as thrust balance feet. The tests on the two EP systems have shown accurate and repeatable results, demonstrating that the balance can be used to characterize different EP systems in the μN–mN thrust range. [ABSTRACT FROM AUTHOR]

The electromagnetic sail, a new space propulsion system without propellant, which propels a spacecraft by deflecting solar wind with an electromagnetic field, is proposed. The thrust mechanism and the performance variations of the electromagnetic sail are analyzed with both particles in cell and magnetohydrodynamic models, and the interaction between the incoming plasma fluid and the electromagnetic field is investigated by tests on the ground experimental system. The simulation results demonstrate that a magnetosphere is created in front of the sail and its area increases with magnetic field intensity; with appropriate dipole moment, the thrust of the electromagnetic sail is higher than a pure magnetic sail or a pure electric sail. In test experiments, the plume structure of a helicon plasma source changes obviously with coil current and metal mesh voltage, which verifies the repulsion of the magnetic field and electric field to the incoming plasma. These results indicate that the combination of the magnetic sail and electric sail can increase the utilization efficiency of solar wind, which means the electromagnetic sail is a feasible propulsion system. [ABSTRACT FROM AUTHOR]

We present a drift kinetic model for the free expansion of a thermal plasma out of a magnetic nozzle. This problem relates to plasma space propulsion systems, natural environments such as the solar wind, and end losses from the expander region of mirror magnetically confined fusion concepts such as the gas dynamic trap. The model incorporates trapped and passing orbit types encountered in the mirror expander geometry and maps to an upstream thermal distribution. This boundary condition and quasineutrality require the generation of an ambipolar potential drop of ∼ 5 T e / e , forming a thermal barrier for the electrons. The model for the electron and ion velocity distributions and fluid moments is confirmed with data from a fully kinetic simulation. Finally, the model is extended to account for a population of fast sloshing ions arising from neutral beam heating within a magnetic mirror, again resulting in good agreement with a corresponding kinetic simulation. [ABSTRACT FROM AUTHOR]

Space propulsion of electroosmotic thrusters (EOTs) with a soft charged nanochannel is investigated considering the Navier slip boundary and constant surface charge density on the walls of slit channels. The soft nanochannel is characterized by a wall-grafted ion-penetrable charged polyelectrolyte layer (PEL). The Poisson–Boltzmann equation is solved to give the electric potential distribution based on the assumption of the Debye–Hückel linearization for the low electric potential. An analytical solution of the electroosmotic velocity through the soft channel is obtained. The thrust, specific impulse, and total input power of EOTs produced by the electroosmotic flow are presented, and then, two significant physical quantities, thruster efficiency and thrust-to-power ratio, are described. It is found that these performance curves strongly depend on the slip length, surface charge density on the walls, drag coefficient, equivalent electric double layer thickness, PEL thickness, and density ratio of the PEL to the electrolyte solution layer. By analyzing and optimizing these design parameters, the simulated EOTs can deliver the thrust from 0 μN to 10 µN as well as the specific impulse from 40 s to 100 s, and the thruster efficiency up to 87.22% is realized. If more thrust control and kinetic energy are needed for different space missions, an array composed of thousands of single EOT emitters is constructed and maintains high thruster efficiency. Moreover, during mission operation, the total potential can be simply varied to optimize the performances of thrusters at any moment. [ABSTRACT FROM AUTHOR]

This paper provides perspectives on recent progress in understanding the physics of devices in which the external magnetic field is applied perpendicular to the discharge current. This configuration generates a strong electric field that acts to accelerate ions. The many applications of this set up include generation of thrust for spacecraft propulsion and separation of species in plasma mass separation devices. These "E × B" plasmas are subject to plasma–wall interaction effects and to various micro- and macroinstabilities. In many devices we also observe the emergence of anomalous transport. This perspective presents the current understanding of the physics of these phenomena and state-of-the-art computational results, identifies critical questions, and suggests directions for future research. [ABSTRACT FROM AUTHOR]

The main emphasis of this paper is on the use of Retarding Potential Analyzers (RPAs) for measuring the electron and ion distribution functions, density, and temperature of space plasmas and the charging level of spacecraft surfaces interacting with the plasmas. Multiple grids of progressively negative potential are used to suppress the secondary electrons produced at the entrance of an RPA. We point out that it is impossible to achieve complete suppression. The secondary electrons collected by the receiver may produce a spurious lump in the distribution measured. Improved designs and space applications for spacecraft charging and ionic liquid ion beam diagnostics are discussed. Spacecraft charging will be very important for planetary explorations, whereas ionic liquid ion beams will be important for spacecraft propulsion in the future. [ABSTRACT FROM AUTHOR]

In negative ion sources, a cusped magnetic field is generated by magnets placed around each aperture of the extraction grid in order to limit the co-extracted electron current. In spite of this suppression magnetic field, the co-extracted electron current is large, on the same order as the negative ion current extracted from the plasma. In this paper, we study the mechanisms of electron extraction from the plasma through a cusped aperture in a simplified situation, in the absence of negative ions, with the help of a three-dimensional Particle-In-Cell Monte Carlo Collisions model. The calculation results show that the electron current extracted from the plasma is small for an infinite slit aperture with a suppressed (cusped) magnetic field and significantly increases in the case of finite slit or circular grid apertures. We find that the E × B electron drift plays an important role in the extraction of electrons through a finite slit grid aperture and that current driven micro instabilities are present in the aperture region. This work is relevant to negative ion sources and micro-ECR neutralizers designed for space propulsion. [ABSTRACT FROM AUTHOR]

There are a number of pressing problems mankind is facing today that could, at least in part, be resolved by space systems. These include capabilities for fast and far-reaching telecommunication, surveying of resources and climate, and sustaining global information networks, to name but a few. Not surprisingly, increasing efforts are now devoted to building a strong near-Earth satellite infrastructure, with plans to extend the sphere of active life to orbital space and, later, to the Moon and Mars if not further. The realization of these aspirations demands novel and more efficient means of propulsion. At present, it is not only the heavy launch systems that are fully reliant on thermodynamic principles for propulsion. Satellites and spacecraft still widely use gas-based thrusters or chemical engines as their primary means of propulsion. Nonetheless, similar to other transportation systems where the use of electrical platforms has expanded rapidly, space propulsion technologies are also experiencing a shift toward electric thrusters that do not feature the many limitations intrinsic to the thermodynamic systems. Most importantly, electric and plasma thrusters have a theoretical capacity to deliver virtually any impulse, the latter being ultimately limited by the speed of light. Rapid progress in the field driven by consolidated efforts from industry and academia has brought all-electric space systems closer to reality, yet there are still obstacles that need addressing before we can take full advantage of this promising family of propulsion technologies. In this paper, we briefly outline the most recent successes in the development of plasma-based space propulsion systems and present our view of future trends, opportunities, and challenges in this rapidly growing field. [ABSTRACT FROM AUTHOR]

Even after 50 years of development, narrowband high-power microwave (HPM) source technologies remain the focus of much research due to intense interest in innovative applications of HPMs in fields such as directed energy, space propulsion, and high-power radar. A few decades ago, the main aim of investigations in this field was to enhance the output power of a single HPM source to tens or hundreds of gigawatts, but this goal has proven difficult due to physical limitations. Therefore, recent research into HPM sources has focused on five main targets: phase locking and power combination, high power efficiency, compact sources with a low or no external magnetic field, high pulse energy, and high-power millimeter-wave generation. Progress made in these aspects of narrowband HPM sources over the last decade is analyzed and summarized in this paper. There is no single type of HPM source capable of excellent performance in all five aspects. Specifically, high pulse energy cannot be achieved together with high power efficiency. The physical difficulties of high power generation in the millimeter wave band are discussed. Semiconductor-based HPM sources and metamaterial (MTM) vacuum electron devices (VEDs) are also commented on here. Semiconductor devices have the advantage of smart frequency agility, but they have low power density and high cost. MTM VEDs have the potential to be high power efficiency HPM sources in the low frequency band. Moreover, problems relating to narrowband HPM source lifetime and stability, which are the important determinants of the real-world applicability of these sources, are also discussed. [ABSTRACT FROM AUTHOR]

This paper presents a technology in microthruster design: the ferroelectric plasma thruster (FEPT). The FEPT utilizes an applied rf electric field to create plasma on the surface of a ferroelectric dielectric. Acceleration of ions from this plasma provides thrust. Advantages of the FEPT include emission of both electrons and ions leading to self-neutralization, creation of plasma, and acceleration of ions with a single power supply, and application of thrust in a short amount of time. We present the concept of the thruster, operational physics, as well as experimental results demonstrating plasma creation and ion acceleration. These results along with plasma spectroscopy allow us to calculate thruster parameters. [ABSTRACT FROM AUTHOR]

Propagation of an ionization front in the beam channel was observed after plasma was generated using a 170 GHz millimeter-wave beam in the atmosphere. The propagation velocity of the ionization front was found to be supersonic when the millimeter-wave power density was greater than 75 kW cm-2. The momentum coupling coefficient Cm, a ratio of the propulsive impulse to the input energy, was measured using conical and cylindrical thruster models. A Cm value greater than 350 N MW-1 was recorded when the ionization front propagated with supersonic velocity. [ABSTRACT FROM AUTHOR]

In this study, a numerical model is developed to simulate the xenon plasma plume from a thermionic hollow cathode employing an orifice plate used for propellant ionization and beam neutralization in an electrostatic space propulsion system. The model uses a detailed fluid model to describe the electrons and a particle-based kinetic approach is used to model the heavy xenon ions and atoms. A number of key assumptions in terms of physical modeling and boundary conditions of the simulations are assessed through direct comparisons with experimental measurements. For two of the three cathode operating conditions considered, good agreement with the measured data is obtained. The third condition appears to lie in a different physical regime where elevated electron and ion temperatures and decreased transport coefficients are required in the simulation to provide agreement between the model and the measured data. © 2004 American Institute of Physics. [ABSTRACT FROM AUTHOR]

Conventional annular Hall thrusters become inefficient when scaled to low power. An alternative approach, a 2.6 cm miniaturized cylindrical Hall thruster with a cusp-type magnetic field distribution, was developed and studied. Its performance was compared to that of a conventional annular thruster of the same dimensions. The cylindrical thruster exhibits discharge characteristics similar to those of the annular thruster, but it has a much higher propellant ionization efficiency. Significantly, a large fraction of multicharged xenon ions might be present in the outgoing ion flux generated by the cylindrical thruster. The operation of the cylindrical thruster is quieter than that of the annular thruster. The characteristic peak in the discharge current fluctuation spectrum at 50-60 kHz appears to be due to ionization instabilities. In the power range 50-300 W, the cylindrical and annular thrusters have comparable efficiencies (15%-32%) and thrusts (2.5-12 mN). For the annular configuration, a voltage less than 200 V was not sufficient to sustain the discharge at low propellant flow rates. The cylindrical thruster can operate at voltages lower than 200 V, which suggests that a cylindrical thruster can be designed to operate at even smaller power. [ABSTRACT FROM AUTHOR]

Charge exchange between xenon ions and xenon atoms is the source of a detrimental low energy plasma in the vicinity of electrostatic spacecraft thrusters. Proper modeling of charge-exchange induced spacecraft interactions requires knowledge of the respective charge-exchange cross sections. Guided-ion beam measurements and semiclassical calculations are presented for xenon atom charge-exchange collisions with Xe[sup +] and Xe[sup 2+] at energies per ion charge ranging from 1 to 300 eV. The present measurements for the symmetric Xe[sup +]+Xe exchange system are in good agreement with several earlier experimental studies and semiclassical calculations based on the most recently computed Xe[sub 2][sup +] interaction potentials. The cross sections are ∼30% higher than predictions by the Rapp and Francis model [D. Rapp and W. E. Francis, J. Chem. Phys. 37, 2631 (1962)]. The present Xe[sup 2+]+Xe symmetric charge exchange measurements are the first to cover the ion energy range from 40 to 600 eV. The cross sections are in good agreement with low-energy drift tube measurements and are significantly lower than previous higher energy measurements. A simple model for symmetric two-electron transfer is proposed that is in good agreement with the present measurements. The onset for the asymmetric charge-exchange process, Xe[sup 2+]+Xe→2Xe[sup +], is observed to be at 10 eV. For this process, a cross section of 2.8±0.9 Å[sup 2] is measured for a Xe[sup 2+] energy of 600 eV. © 2002 American Institute of Physics. [ABSTRACT FROM AUTHOR]

Ionic liquid ion source (ILIS) is a promising ion source, which can be applied to space propulsion, microfabrication, and surface modification. Fabrication of high-quality ILIS emitters is one of the key technologies for the application of ILIS. A new method is proposed for the fabrication of porous emitters with a designed shape. This method uses wire electrical discharge machining (WEDM) combined with electrochemical etching, and the porous emitter is fabricated by two steps. First, the porous metal is machined by WEDM to get the external geometry of the emitter. Then, electrochemical etching is employed to remove the recast layer. A series of experiments has been conducted to find the appropriate machining parameters. Experiments reveal that sharp porous emitter tips for the ILIS can be fabricated by WEDM combined with electrochemical etching at 5 V etching voltage. Moreover, the apex curvature radius of the emitter is controllable by adjusting the etching time. It is found that the apex curvature radius varies from 4.5 μm to 18.4 μm when increasing etching time from 40 s to 120 s at 5 V etching voltage. Those emitters have been applied to ILIS tests, and their I-V characteristics are investigated. Furthermore, this method has been used to machine dense fields of emitters. A 1 cm2 emitter array chip integrated with 676 emitters has been successfully machined, and the I-V characteristic curve of the emitter array chip is also achieved. [ABSTRACT FROM AUTHOR]

A steady-state plasma flow exiting through an expanding magnetic field is studied by means of 2D particle in cell numerical simulations. The effects of change in the localized plasma source region dimension and the associated plasma transit length in an upstream, uniform-magnetic-field region were examined, simulating cases with different axial lengths of the dielectric plasma source region, distinct from the location of physical expansion. Axial potential profiles at various radial locations show development of a stepwise axial potential drop, producing plasma (ion) acceleration in the corresponding regions. For a narrow source region with a long pre-expansion flow region, nonmonotonic potential variation is recovered where ion phase-space scatters show the trapped region and signatures of chaotic ion trajectories and possible pre-expansion ion heating. Considering a relevance of the studied flow equilibria to the thrust generation schemes for space propulsion, a formal estimate of thrust values associated with the plasma outflow is also done for the cases simulated. [ABSTRACT FROM AUTHOR]

In spite of its high cost, xenon gas is known as both the most efficient and commonly used propellant for plasma thrusters in space technologies. Argon, a gas by far less costly, is widely used in other technologies, but a much lower efficiency of ionization, as obtained for example in closed electron drift thrusters, prevents its use in R&D programs and development of space thrusters. This paper shows that a drastic increase in argon ionization can be obtained in a low power thruster when only a few percent of xenon are added in the argon flow. Besides the increase in the ion beam current in the plume generated by the thruster, a net increase in the ion kinetic energy is observed. These two features are of interest in terms of thrust efficiency. These results, obtained for a small size closed electron drift thruster, could be even more spectacular for higher power devices, suggesting further investigations for space propulsion and/or ion source applications. [ABSTRACT FROM AUTHOR]

Space propulsion systems use screen mesh devices as filters to block contaminants and as propellant management devices to settle the propellants. The bubble point pressure indicates the basic capillary performance for liquid acquisition of screen meshes. Actively controlling the bubble point pressure can result in flexible and efficient operation of the propulsion systems. High-performance cryogenic propellants, such as liquid hydrogen and oxygen, exhibit magnetic properties. Therefore, a method to actively control the bubble point pressure of cryogenic propellants by applying a magnetic field is proposed in this study. The magnetic pressures affect the pressure balance around the gas–liquid interface separated by the screen mesh, which can thereby control the bubble point pressure. To demonstrate the concept and theoretical basis, a bubble point experiment is conducted using a ferrofluid and solenoid. This experiment proves that the magnetic field actively controls the bubble point pressure and performs both suppression and enhancement of the liquid acquisition performance of the screen mesh. The theory related to magnetic pressures is observed to successfully predict the experimental results. The feasibility of the active control of the bubble point pressure of liquid oxygen is discussed based on the validated theory, and two applications of this technique in cryogenic propulsion systems are depicted. [ABSTRACT FROM AUTHOR]

(2) The title should read as "The generalized plasma focus problem and its application to space propulsion." REFERENCES 1 S. K. H. Auluck, " Generalized plasma focus problem and its application to space propulsion", Phys. Erratum: "The generalized plasma focus problem and its application to space propulsion" [Phys. [Extracted from the article]

Most state-of-the-art electric space propulsion systems such as gridded and Hall effect thrusters use xenon as the propellant gas. However, xenon is very rare, expensive to produce, and used in a number of competing industrial applications. Alternatives to xenon are currently being investigated, and iodine has emerged as a potential candidate. Its lower cost and larger availability, its solid state at standard temperature and pressure, its low vapour pressure and its low ionization potential make it an attractive option. In this work, we compare the performances of a gridded ion thruster operating separately with iodine and xenon, under otherwise identical conditions using a global model. The thruster discharge properties such as neutral, ion, and electron densities and electron temperature are calculated, as well as the thruster performance parameters such as thrust, specific impulse, and system efficiencies. For similar operating conditions, representative of realistic thrusters, the model predicts similar thrust levels and performances for both iodine and xenon. The thruster efficiency is however slightly higher for iodine compared with xenon, due to its lower ionization potential. This demonstrates that iodine could be a viable alternative propellant for gridded plasma thrusters. [ABSTRACT FROM AUTHOR]