Search

Scaling methods are a straightforward way to design new thrusters at different power levels. Based on the general equations, this work derives the scaling rules from those equations and presents a model for thruster with anode layer (TAL) and stationary plasma thrusters (SPTs) that encompass the scaling methods described until now. With the exception of the magnetic field, the scaling rules derived can be used for both TAL and SPT. In addition, this model is used to study the impact of the different scaling methods in the performance of the thruster. In particular, how the total thruster efficiency is affected through the current efficiency and the influence of this efficiency loss in the thrust. Accordingly, using the rules derived from the general equations for the magnetic field, the possible conflicts with the electron confinement given by the Larmor Radius for both SPT and TAL thrusters are studied. Finally, it is shown how the selected scaling method strongly affects the expected lifetime of the thruster. In this aspect, the strict scaling results to be the most rigorous method, whereas the length extension method allows for further improvement of the lifetime. [ABSTRACT FROM AUTHOR]

When designing spacecraft with electric propulsion (EP) devices, it is important to assess spacecraft integration to ensure that the important components are not subject to significant sputtering by high-energy ions. In addition to the EP device and its plume, surface charging of spacecraft has to be modeled properly, as surface potential can directly affect the sputtering rate. Three surface charging models are incorporated into the spacecraft module of the numerical simulation framework–Thermophysics Universal Research Framework (TURF), and these include: 1) dielectric; 2) conductive; and 3) charge propagation models. The charge propagation model has been upgraded to solve the surface charge distribution implicitly, allowing a wide range of electrical conductivity values without causing the simulation to become unstable. Each of the charging models is verified against a simple problem where an analytical solution can be determined. Then, the coupling of the surface charging model and a hybrid particle/fluid model is tested in a more complex problem, where the floating potential on a sphere immersed in plasma is to be obtained. Finally, the surface charging model in an EP plume simulation is demonstrated. These problems are multiscale in that the charging model has to resolve an electron timescale (i.e., plasma oscillation) while the particle time step has to be orders of magnitude larger than the electron timescale in order to maintain a long enough sampling window for the ion current to effectively reduce the statistical noise. Therefore, two separate time steps are introduced for a stable convergence of the coupled models. [ABSTRACT FROM AUTHOR]

The acceleration of rotating plasma flow in crossed magnetic fields produced momentum transfer between the macroscopic degrees of freedom for a plasma flow was investigated. Results implied a concept for acceleration of plasma flow containing the charged macroscopic particles. Here, the case is treated when the plasma flow consists of electrons, protons, and heavy, multiply charged negative dust particles only. The analysis shows that these charged macroscopic particles have been trapped and, then, accelerated in the main plasma flow. As a result, such complex plasma flow can be accelerated in one axial direction to increase the thrust. It suggests that such complex plasma flow produced, for example, from cosmic dust medium may be used as a propellant and a work body for plasma thrusters. [ABSTRACT FROM AUTHOR]

Abstract Ignition process, as the initiation of the entire discharge, plays an important role in ablative pulsed plasma thrusters. While spark plug exactly how initiate discharge is achieved is still under review. This study did some experiments with two kinds of propellant surfaces (normal or inclined) and without propellant to explain the ignition process. The experimental results showed: when the thruster discharge without propellant, it is essentially a surface flashover process on ceramics; when the propellant was loaded, the main discharge occurs after the initial conductive path composed of electrons emitted by spark plug forming. This study provides a reference for the high performance pulsed plasma thrusters. [ABSTRACT FROM AUTHOR]

This paper presents calibration devices and methods for the measurement of electric thruster performance parameters using a seesaw-type thrust stand to measure the mass loss of solid propellant in a vacuum. In previous studies, impact hammers and electrostatic combs have been manufactured for the calibration of the thrust and impulse using seesaw-type thrust stands. However, these conventional devices rely on self-calibration, which means that the input delivered by the device in unknown, and must undergo a calibration process themselves. In this paper, the manufactured calibration devices successfully reproduced known impulses, thrusts, and mass losses in a vacuum. By reproducing known inputs based on known masses, the proposed calibration devices can omit the conventionally required self-calibration process. The calibration results showed linear relations between outputs and known inputs and agreed with the theoretical values to within an error of 10%. Additionally, the uncertainties of all known inputs were less than 1.5%. On the basis of these results, the average thrust, impulse, and mass loss were measured using a calibrated thrust stand for the first time. The cumulative impulses obtained from the measured impulse and average thrust agreed with each other to within an error of 5%. The error of the measured mass loss per 1000 shots with respect to the actual mass loss measured using an electronic balance ranged from 1% to 17%. [ABSTRACT FROM AUTHOR]

Low power capillary discharge based pulsed plasma thrusters (CDPPTs) are electrothermally dominated thrusters and they have aroused renewed interest in the investigation and enhancement of the basic performance for the application of micro/nano satellites recently. Research on the ignition mechanism of a CDPPT has been conducted to provide insight into the optimization of the structure design and promotion of the lifetime performance. It has been found that the electrical parameters, the geometry parameters, and the cumulative effect of discharge jointly determine the discharge ignition characteristics. A single ignition process is divided into the breakdown of igniter and the development of the main discharge, while the results show that the jitter of the ignition delay time is mainly introduced from the former. Shorter ignition delay time and lower jitter can be obtained with a higher ignition energy and main charging voltage or a shorter and narrower cavity, which is positively correlated to the electric field distribution along the propellant surface. Moreover, with long duration experiments, it reveals that the surface deposits and morphology of propellant and igniter are the dominant factors that cause the dispersity of the ignition effect and main discharge characteristics. [ABSTRACT FROM AUTHOR]

We examine the creation and design of a plasma thruster based on the converting of radial and circumferential momentum of the plasma flow into the axial direction. The acceleration effect is formed by the azimuthal electric field which comes about as a variation of an external, axial magnetic flux. Simultaneously, the azimuthal electron and ion flows interact with the radial magnetic field via the Lorentz force. Owing to this process, the transformation of azimuthal momentum transfers into axial momentum. As a result, the electron and ion flows are accelerated in one axial direction. Estimations specifying the admissible range of parameters of this device and the plasma flow are derived. [ABSTRACT FROM AUTHOR]

Fully kinetic plasma simulations, in which ions and electrons are represented by computational particles, are attractive for modeling electron-ion coupling in spacecraft electric propulsion devices. While conceptually simple, fully kinetic simulations are susceptible to a numerical instability due to a finite computational domain. The widely utilized workaround is to artificially expand the domain and stop the simulation long before the plume reaches the mesh boundaries. The simulation thus never reaches the steady state, and massive computational resources are required. This paper describes a new boundary treatment for fully kinetic studies that allows reaching steady state on a relatively small domain. This approach is demonstrated by first modeling a preneutralized beam. The fully kinetic results are compared with a hybrid model with fluidlike electrons. A more complex setup containing a centrally mounted cathode is investigated next. Grid independence is verified by comparing results on two different sized domains. [ABSTRACT FROM AUTHOR]

In the previous investigations, an intermode exchange was used to demonstrate the transfer of energy and momentum into using different kinematic degrees of freedom. This paper examines using a hydrodynamic approach using a cold, neutral flow in a rotating cylindrical plasma column using crossed magnetic fields. Results show that the twirling plasma flow in a cylindrical vortex can be accelerated in an axial direction, thereby resulting an energy/momentum transfer in the axial direction. These results are analogous to the creation of a plasma thruster using this effect. [ABSTRACT FROM PUBLISHER]

Methods, algorithms and means of few projections spectral and tomographic diagnostics of parameters of plasma flows, such as space distribution of local temperatures, concentration, etc. of small-size plasma ejectors (engines of orientation, low thrust) are described. [ABSTRACT FROM AUTHOR]

A small information satellite which may be placed on an orbit with altitude from 600 up to 1000 km is considered. The satellite attitude and orbit control system contains a strap-down inertial navigation system, cluster of four reaction wheels, magnetic driver and a correcting engine unit with eight electro-reaction engines. Problems on the design of algorithms for control of the satellite spatial motion on sun-synchronous orbit are studied. [ABSTRACT FROM AUTHOR]

The miniature ion-plasmatic engine for Cube Sate nanosatellites is described. The working body is the mercury ions received from metal mercury by radiation from its cesium source. The offered construction allows NS trajectory parameters to be changed, to compensate oscillating and rotational motions of the device. Examples of calculation of the IPE some parameters are given. [ABSTRACT FROM AUTHOR]

Plasma engines are used for space propulsion as an alternative to chemical thrusters. Due to the high exhaust velocity of the propellant, they are more efficient for long-distance interplanetary space missions than their conventional counterparts. An advanced laboratory of plasma space propulsion (PlaNS) at the Institute of Plasma Physics and Laser Microfusion (IPPLM) specializes in designing and testing various electric propulsion devices. Inside of a special vacuum chamber with three performance pumps, an environment similar to the one that prevails in space is created. An innovative Micro Pulsed Plasma Thruster (LμPPT) with liquid propellant was built at the laboratory. Now it is used to test the second prototype of Hall effect thruster (HET) operating on krypton propellant. Meantime, an improved prototype of krypton Hall thruster is constructed. [ABSTRACT FROM AUTHOR]

We have developed a unique experimental device of so-called double plasma gun, which consists of two magnetized coaxial plasma gun (MCPG) devices, in order to clarify effects of vapor shielding on material erosion due to transient events in magnetically confined fusion devices. Two ELM-like pulsed plasmas produced by the two MCPG devices were injected into a target chamber with a variable time difference. For generating ablated plasmas in front of a target material, an aluminum foil sample in the target chamber was exposed to a pulsed plasma produced by the 1st MCPG device. The 2nd pulsed plasma was produced with a time delay of 70 μs. It was found that a surface absorbed energy measured by a calorimeter was reduced to ∼66% of that without the Al foil sample. Thus, the reduction of the incoming plasma energy by the vapor shielding effect was successfully demonstrated in the present experiment. [ABSTRACT FROM AUTHOR]

The ignition of LOVA gun propellants at low loading densities is of interest for future gun systems with insensitive modular charges. The ignition system in conventional propellant systems is traditionally optimized for high loading densities. There is a need to be able to control the ignition for the envelop of the system, to reduce unburnt propellant at low loading densities and, at the same time, avoid pressure waves at high loading densities. Plasma ignition is interesting for this application, with the possibility to vary ignition energy and with the plasma plume penetrating deep into the propellant bed to achieve uniform ignition. This paper presents results from plasma ignition experiments in a 45-mm gun system, using NL008, an RDX-based CAB-LOVA propellant. The plasma igniter is a capillary jet type, driven by a pulse-forming network. A zero-dimensional model for the plasma igniter has been developed and will be discussed in the context of the experiments. [ABSTRACT FROM PUBLISHER]

Impulse bits produced by conical theta-pinch inductive pulsed plasma thrusters possessing cone angles of 20°, 38°, and 60°, were quantified for 500-J/pulse operation by direct measurement using a hanging pendulum thrust stand. All three cone angles were tested in single-pulse mode, with the 38° model producing the highest impulse bits at roughly 1-mN-s operating on both argon and xenon propellants. A capacitor charging system, assembled to support repetitively pulsed thruster operation, permitted testing of the 38° thruster at a repetition rate of 5 Hz at power levels of 0.9, 1.6, and 2.5 kW. For similar conditions, the average thrust measured during repetitive-pulse operation exceeded the value obtained when the single-pulse impulse bit is multiplied by the repetition rate, suggesting that a greater impulse bit per pulse was produced when operating in the repetitive-pulse mode. [ABSTRACT FROM PUBLISHER]

In this paper, the design and performances of a pulsed plasma thruster (PPT) for nanosatellite applications will be presented. The breadboard model PPT presented in this paper will be a part of six PPTs propulsion system designed to provide attitude and translational control for a 20-kg nanosatellite for a total delta-V of 40 m/s. The thruster performances have been characterized in terms of electrical parameters, mass bit, impulse bit, and specific impulse. The thruster was found to be compliant with the mission requirement. Moreover, preliminary electromagnetic noise interference measurements have been performed. The spark plug discharge was found to be the main source of noise as already found by previous authors. From the collected data, it can be inferred that the noise is mostly radiated. [ABSTRACT FROM PUBLISHER]

Understanding the physics involved in plasma detachment from magnetic nozzles is well theorized, but lacking in large scale experimental support. We have undertaken an experiment using the 150-m ^3 variable specific impulse magnetoplasma rocket test facility and VX-200 thruster seeking evidence that detachment occurs and an understanding of the physical processes involved. It was found that the plasma jet in this experiment does indeed detach from the applied magnetic nozzle (peak field \sim 2 T) in a two part process. The first part involves the ions beginning to deviate from the nozzle field 0.8-m downstream of the nozzle throat. This separation location is consistent with a loss of adiabaticity where the ratio of the ion Larmor radius to the magnetic field scale length ( r_{Li}|\nabla B|/B ) becomes of order unity and conservation of the magnetic moment breaks down. Downstream of this separation region, the dynamics of the unmagnetized ions and magnetized electrons, along with the ion momentum, affect the plume trajectory. The second part of the process involves the formation of plasma turbulence in the form of high-frequency electric fields. The ion and electron responses to these electric fields depend upon ion momentum, magnetic field line curvature, magnetic field strength, angle between the particle trajectories, and the effective momentum transfer time. In stronger magnetic field regions of the nozzle, the detached ion trajectories are affected such that the unmagnetized ions begin to flare radially outward. Further downstream as the magnetic field weakens, for higher ion momentum and along the edge of the plume, the fluctuating electric field enables anomalous cross-field electron transport to become more dominant. This cross-field transport occurs until the electric fields dissipate \sim 2$ -m downstream of the nozzle throat and the ion trajectories become ballistic. This transition to ballistic flow correlates well with the sub-to-super Alfvénic flow transition ( \beta _{k} ). There was no significant change observed to the applied magnetic field. [ABSTRACT FROM PUBLISHER]

A thrust balance characterization of a low powered Quad Confinement Thruster is presented for high levels of propellant flow. The nominal flow rate for this device is between 1 and 2 sccm of xenon propellant. This paper extends the operating range, and investigates the performance at two high flow conditions of 10 and 20 sccm. Power is varied incrementally between 20 and 200 W in order to characterize the performance versus power trends of the device. It was found that for these high flow regimes the propellant is underutilized, and a proportion of the increased thrust can likely be attributed to a hot gas expansion of the neutral xenon rather than the generation of additional accelerated ions. The thrust was increased from 1 (nominal) to 3.3 mN at 200 W of input power for the 20 sccm condition. However, the performance penalty in terms of the specific impulse was considerable. The specific impulse under these conditions dropped below 200 s, where the nominal condition is 1000 s. A compromise between increased thrust and decreased performance was found at 10 sccm of flow: 3 mN of thrust at 300 s specific impulse. [ABSTRACT FROM PUBLISHER]

The field of application of Hall effect thrusters (HETs) is constantly expanding toward increased power and specific impulse and also toward reduced power. The modern level of plasma simulations does not allow accurate prediction of a thruster performance in advance. Therefore, the methods of scaling play an important role in the creation of new thrusters with desired characteristics. This paper describes a scaling model of HETs based on an analytical assessment of the anode mass utilization efficiency and the available experimental data. Empirical coefficients of the model are found using an extensive database containing published test results of many thrusters. The obtained expressions allow for predicting the performance of HETs for various kinds of propellant when the discharge power and voltage vary over a wide range. [ABSTRACT FROM PUBLISHER]

Physical processes determining the excitation of RF electromagnetic fields in a plasma column in a magnetic field are analyzed. The Alfvén resonance plays an important role at frequencies close to the ion cyclotron frequency. It leads to the enhancement of the RF electric field and transformation of Alfvén oscillations with a predominantly transverse polarization of the electric field into lower hybrid ones, which have a significant longitudinal component of the electric field. Lower hybrid oscillations efficiently interact with electrons causing their heating. Difficulties in the implementation of ion cyclotron resonance heating by the magnetic beach method are outlined. The processes considered in this work can be important for the VASIMR plasma engine. [ABSTRACT FROM AUTHOR]

A particle-in-cell Monte Carlo collision model is developed to explore dominant collisional effects on high-velocity xenon ions incident to a quiescent xenon gas at low neutral pressures. The range of neutral pressure and collisionality examined are applicable for electric propulsion as well as plasma processing devices; therefore, the computational technique described herein can be applied to more complex simulations of those devices. Momentum and resonant charge-exchange collisions between ions and background neutrals are implemented using two different models, classical scattering with spin-orbit free potential and variable-hard-sphere model, depending on the incident particle energy. The primary and charge-exchange ions are tracked separately, and their trajectories within a well-defined “Test Cell” domain are determined. Predicted electrode currents as a function of the Test Cell pressure are compared with electrode currents measured in an ion gun experiment. The simulation results agree well with the experiment up to a Test Cell pressure corresponding to a mean free path of the Test Cell length and then start to deviate with increasing collisionality at higher pressures. This discrepancy at higher pressures is likely due to the increasing influence of secondary electrons emitted from electrodes due to the high-velocity impacts of heavy species (i.e., beam ions and fast neutrals created by charge-exchange interaction) at the electrode surfaces. [ABSTRACT FROM PUBLISHER]

A model of pulsed inductive plasma thrusters consisting of a set of coupled circuit equations and a 1-D momentum equation has been used to study the effects of adding a second parallel capacitor into the system. The equations were nondimensionalized, permitting the recovery of several already known scaling parameters and leading to the identification of a parameter that is unique to the particular topology studied. The current rise rate through the inductive acceleration coil was used as a proxy measurement of the effectiveness of inductive propellant ionization since higher rise rates produce stronger, potentially better ionizing electric fields at the coil face. Contour plots representing thruster performance (exhaust velocity and efficiency) and current rise rate in the coil were generated numerically as a function of the scaling parameters. The analysis reveals that, when the value of the second capacitor is much less than that of the first capacitor, the performance of the two-capacitor system approaches that of the single-capacitor system. In addition, as the second capacitor is decreased in value, the current rise rate can grow to be twice as great as the rise rate attained in the single-capacitor case. [ABSTRACT FROM AUTHOR]

Spacecraft charging codes require accurate models of electron yields as a function of accumulated charge to correctly predict the charge buildup on spacecraft. The accumulated charge creates equilibrium surface potentials on spacecraft resulting from interactions with the space plasma environment. There is, however, a complex relation between these emission properties and the charge built up in spacecraft insulators. This paper focuses on different methods appropriate to determine the fundamental electronic material property of the total electron yield as the materials accumulate charge. Three methods for determining the uncharged total yield are presented: 1) The dc continuous beam method is a relatively easy and accurate method appropriate for conductors and semiconductors with maximum total electron yield \sigma\max < 2 and resistivity \rho < 10^17 \ \Omega\cdot\cm; 2) the pulsed-yield method seeks to minimize the effects of charging and is applicable to materials with \sigma\max < 4 and \rho up to >10^24 \ \Omega\cdot\cm; and 3) the yield decay method is a very difficult and time-consuming technique that uses a combination of measurement and modeling to investigate the most difficult materials with \sigma\max > 4 and \rho up to <10^24 \ \Omega\cdot\cm. Data for high-purity polycrystalline Au, Kapton HN and CP1 polyimides, and polycrystalline aluminum oxide ceramic are presented. These data demonstrate the relative strengths and weaknesses of each method but more importantly show that the methods described herein are capable of reliably measuring the total electron yield of almost any spacecraft material. [ABSTRACT FROM PUBLISHER]

Abstract-A compact high-power-density helicon plasma source operating on abundant medium atomic weight gases is under examination for a potential electric propulsion application. A high specific impulse, which is easily varied in the range of ∼4-2 ks for Ar and ∼2-4 ks for N2 propellant, was achieved. A wellcollimated plasma plume with emittance ∼10° is produced in a single-magnet configuration. Analysis of the data obtained using in situ and noninvasive spectral diagnostics indicates almost complete gaseous propellant ionization and thrust efficiency ηT ∼20% with the lowest net ionization cost of ∼(80 V)-1 per ion. Direct weighing has yielded a thrusting force of T ∼ 10 mN. [ABSTRACT FROM AUTHOR]

The properties of high-frequency (HF) instabilities (f > 1 MHz) in Hall thrusters are reviewed on the basis of a large experimental data set previously obtained with the use of antennas, probes, and magnetic coils. The phenomenological representations of their physical sources are discussed. The development of these HF instabilities results in anomalous electron transport, and the corresponding transport coefficients are evaluated from these experimental data. [ABSTRACT FROM AUTHOR]

A particle-in-cell code has been developed for simulations of plasmas of Hall thruster discharges. The simulated system is a plasma slab bounded by dielectric walls with secondary electron emission. An external, accelerating electric field directed parallel to the walls and an external magnetic field directed normal to the walls are applied. The strongly anisotropic non-Maxwellian electron velocity distribution function is obtained in simulations. The average energy of electron motion parallel to the walls is defined by collisional heating in the accelerating electric field. This energy is much higher than the average energy of electron motion normal to the walls, which is determined by the energy of electrons produced in ionization and by scattering of electrons by neutral atoms. The electron distribution function for velocity components normal to the walls is depleted for energies above the near-wall plasma potential. The effects of Coulomb collisions on the electron velocity distribution function and electron wall losses are studied. [ABSTRACT FROM AUTHOR]

Ion emission from ferroelectric plasma sources driven by radio frequency (RF) applied voltage is studied. An experimental investigation of particle emission from lead zirconate titanate ferroelectric ceramics driven by bursts of seven cycles of RF voltage at 248 kHz has revealed significant ion current emission. Measured electron to ion peak current ratios ranged from 11 to 19, which is much lower than the expected thermal current ratio. These results indicate that the ions are being preferentially accelerated. An analysis of the ponderomotive force on the ions and electrons reveals that it may be responsible for acceleration of the ions. The ferroelectric plasma is examined as a possible micropropulsion thruster. Performance estimates based on the ponderomotive acceleration calculations predict specific impulse of 2000-3000 s, thrust of 0.1-1 mN, and specific power of 15 W/mN. These estimates are compared to existing micropropulsion concepts, revealing the attractiveness of ferroelectric plasma sources. [ABSTRACT FROM AUTHOR]