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The Pulsed Inductive Thruster (PIT) has the advantages of repeatable startup, no corruption and in-situ propellant feed. To study the flow expansion and circuit characteristics of PITs, the circuit-fluid model is developed, and the high temperature thermodynamic and transport models are combined with the circuit-fluid model to predict the critical plasma parameters. The flow fields of initial mass of 2–8 mg and charge voltages of 10–14 kV are simulated. Comparison of the flow fields of argon and helium propellants suggests that, the flow field structures are similar. Slight differences exist on the magnitude of the density and magnetic field, caused by larger velocity in lighter atom case and difference on the ionization gap between adjacent ionization levels. Analysis of the circuit characteristics by the two-dimensional results indicates that the ratio of coil inductance to circuit inductance affects both the rise rate and phase of the plasma current, the larger the ratio, the greater the rise rate and the better the following characteristic. The calculations show that the magnetic energy obtained within the decoupling distance determines the overall performance the thruster can be obtained; self-induced field maintained by the thermal motion after the main pulse leads to the long attenuation process and difference on the total impulse when the angle of conical pylon is varied under constant coil dimension.

The pulsed inductive thruster is characterized of no electrode corruption and wide propellant choice. To give insight into the propulsion mechanism of small scale thruster at different propellant mass (m) and energy (E) levels, the transient MagnetoHydroDynamics (MHD) method, completed by high temperature thermodynamic and transport, and plasma electrical models, is developed to study argon plasma response under the excitation of current of high rise rate. By calculating the two-dimensional expansion properties of the thruster with conical pylon, the simulations find that the main energy deposition occurs during the initial pulse rise stage, and the energy density of Joule heat is two magnitudes higher than the deposition in the down side. At propellant mass of 2 mg, average axial velocity of the current sheet increases from about 15 km/s at 750 J to about 21 km/s at 1470 J within the decoupling distance. The velocity variation synchronizes with the pulsed rise in the initial. The monotonically decrease of the temperature along axis results in the growth of low ionization level ions and reducing of high levels. The current sheet maintains the structure formed during the initial pulse rise when moving beyond the decoupling distance. Besides the change in forward velocity, the main difference is the dimension compared with that in the first half period, caused by thermal conduction and particle diffusion. The variations of total impulse It in the range of m from 2 mg to 8 mg and E from 750 J to 1470 J show that It is proportional to m when E is determined. Keywords: Expansion properties, Flow field, MagnetoHydroDynamics (MHD) method, Pulsed inductive plasma, Total impulse

The pulsed inductive thruster accelerates the propellant by the repulsion between inductive coil and current sheet. To accurately investigate the acceleration characteristics in the first half period of pulsed inductive discharge and the energy needed to generate effective impulse, an unsteady magnetohydrodynamics model is developed, in which the coil-plasma boundary condition is improved by plasma electrical model, and the electrical conductivity is calculated using gas kinetic method. The analysis of plasma electrical characteristics shows confining of particles in the beginning and acceleration of the current sheet after ionization process is completed, leaving behind the low-density residual plasma, with negligible contribution to the total impulse. The impulse at high voltage decreases monotonically after peak value is reached, showing effective impulse generation in the first half period, especially before the decoupling distance is reached.

A propulsion concept relying on the direct steady-state acceleration of a plasma by an inductive wave-launching antenna is presented. By operating inductively in steady state, a direct wave-drive t...

A 150T pulsed magnetic field generator device which can generate a 150T pulsed magnetic field by a single-turn coil with a 3 mm diameter was designed to study the coupling of laser-produced plasma and strong magnetic fields is described. An 880 kA current pulse is discharged into the coil by a pulsed-power system composed of 20 capacitors of 5 μF, electric-trigger switches, and low-impedance strip transmission lines. The structure of the generator was optimized based on the results of electromagnetic analysis. A set of small pulsed discharge system with a 50 kV/5 μF capacitor bank was built to research synchronous trigger and the current curve. The circuit verification experiment was proceeded and the result was revealed to be consistent with theoretical analysis.

A 200 W No-Wall-Losses Hall Thruster (NWLHT-200 W) is designed and processed to verify the technology of pushing down magnetic field with two permanent magnetic rings. To create a magnetic field, NWLHT-200 W uses two permanent magnetic rings (inner and outer) in the absence of magnetic screen or magnetic component. The anode is at the internal magnetic separatrix position, and the thruster shell is hollow to enhance the heat dissipation of ceramics. The magnetic field strength at the channel outlet is 90% of the maximum magnetic field. In this study, the experimental results concerning the thrust, discharge current, specific impulse, and efficiency are presented and examined. Our experiments show that “no erosive discharge” of wall is achieved within the range of 120–460 W; the maximum efficiency of the anode may reach 49%. The thruster designed can work stably for a long time, without any auxiliary heat dissipation equipment (heat pipe or radiator), which significantly prolongs the life of Hall thrusters.

Electric propulsion presents an excellent alternative to conventional chemical counterparts as higher exhaust velocities from a plasma-based thruster allow for more efficient mass utilization and expanded space mission capabilities. Understanding the detailed time and space dependant physical phenomena within the main thruster could potentially enhance the capabilities of the electric thruster. Hence, in the present work a two-stage Hall thruster is investigated and the potential feasibility of coupling the two stages to produce the ion-beam is evaluated. The ionization and acceleration events are investigated via multiphysics simulations. The key trends ranging from the discharge inception, amplification and drift to formation of the ion beam are discussed and characterized.

A performance analysis for a pulsed plasma thruster(PPT) has been conducted to predict the thrust and current change. Two models were implemented - a one-dimensional lumped circuit analysis model and the Teflon ablation model provided by Michael Keidar. The circuit model incorporating resistance and inductance models was adapted to predict the magnitude of the discharge current. Numerical simulations like current discharge rates with different voltages were reasonably well compared with experimental data. The effects of Teflon ablation on thruster characteristics were investigated.

The characterization of a miniaturized ionic liquid electrospray thruster for nanosatellite applications is presented. The thruster investigated features an emitter array of 480 emitter tips per square centimeter and a 1 cm[superscript 3] propellant tank with an entirely passive propellant supply; it is operated at a power level of less than 0.15 W. The paper presents energy- and mass-resolving beam spectroscopy of the packaged thruster system, as well as two independent thrust measurements. This allows derivation of thruster performance parameters under realistic firing conditions, including individual thruster efficiency contributions, specific impulse, and thrust. The total thruster efficiencies of 36%; specific impulse of ∼760 s, including all losses; and thrust of 11–12.5 μN are presented at emission currents of 150 μA for a device of ∼1 cm[superscript 2]. The current emission data without current decay of ∼90 h are presented with a maximum of 172 h., United States. National Aeronautics and Space Administration (Contract NNL13AA12C)

The back-sputtering deposition on thruster surface caused by ion bombardment on chamber wall material affects the performance of thrusters during the ground based electric propulsion endurance tests. In order to decrease the back-sputtering deposition, most of vacuum chambers applied in electric propulsion experiments are equipped with anti-sputtering targets. In this paper, a three-dimensional model of plume experimental system (PES) including double layer anti-sputtering target is established. Simulation cases are made to simulate the plasma environment and sputtering effects when an ion thruster is working. The particle in cell (PIC) method and direct simulation Monte Carlo (DSMC) method is used to calculate the velocity and position of particles. Yamamura's model is used to simulate the sputtering process. The distribution of sputtered anti-sputtering target material is presented. The results show that the double layer anti-sputtering target can significantly reduce the deposition on thruster surface. The back-sputtering deposition rates on thruster exit surface for different cases are compared. The chevrons on the secondary target are rearranged to improve its performance. The position of secondary target has relation with the ion beam divergence angle, and the radius of the vacuum chamber. The back-sputtering deposition rate is lower when the secondary target covers the entire ion beam.

The high amplitude pulse current during Hall thruster start-up is an important source of electromagnetic interference, which impacts the safety and reliability of the propulsion system. The characteristics of the pulse current, including the pulse current's integrated area, peak value, and transient time, are studied experimentally and theoretically under different Hall thruster discharge parameters. Results show that the pulse current's integrated area is determined by the atom density in the channel before Hall thruster start-up; the effects of coil current on the peak value and transient time of the pulse current are negligible. As the discharge voltage and mass flow rate increase, the pulse current peak value increases, the pulse current duration is inversely proportional to the discharge voltage and directly proportional to the mass flow rate.

Development of electric propulsion systems (EPS) with high efficiency, reliability and long lifetime is required for solution of present-day space challenges. Ion thrusters (IT) are among the most promising types of EPS. This work is dedicated to numerical simulation of discharge chamber plasma. A 2D 3 V kinetic non-stationary particle-in-cell model has been developed. It includes several new techniques. Maintenance of quasi-neutrality at the cathode border allows the value of discharge current depending on the dynamic plasma parameters to be obtained. In addition, the model includes a technique for modelling operation of a power supply. It expects discharge current control by discharge voltage variation. Combination of these two features allows us to consider plasma discharge in the DC as a dynamic self-consistent system. Several simulations of IT-50 ion thruster operation modes were carried out using the present model. Simulation results and experimental data show sufficient agreement.

A basic scheme of an ablation pulsed plasma thruster with rail geometry and a side feed of the propellant bar is considered. For such geometry, a theoretical description of the initial stage of the vacuum discharge, which would serve to explain the experimental results, does not exist. It was demonstrated experimentally that the configuration of the discharge channel depends essentially on the initial stage of discharge. A model of the motion of charged particles in the inhomogeneous electric and magnetic fields was developed. Due to the low plasma density at the initial stage of discharge, the model of the charged particles motion in electric and magnetic fields is described well by differential equations.

This paper presents the results of plasma parameters probe studies in the discharge chamber of low power laboratory ion thruster. The measurement procedure based on the use of triple Langmuir probe placed into the discharge chamber's plasma. During the plasma measurements probe characteristics were collected automatically by data acquisition software created in LabVIEW development environment. In the current paper we discuss the results of the experiment: electron temperature T e distribution and electron number density n e distribution in the plasma in the hemispherical discharge chamber of the low power laboratory ion thruster. The distributions were obtained without applying an accelerating potential in the ion-optical system for different values of the propellant (xenon) mass flow. Results obtained are in good agreement with the similar experimental data presented in publications of other authors.

The article is devoted to the research of low-power (300 W) radio-frequency ion thruster designed at the Moscow Aviation Institute. The main results of experimental research of the thruster using the testfacility power supplies and the power processing unit of their own design are presented. The dependence of the working fluid ionization cost on its mass flow rate at the constant ion beam current was investigated experimentally. The influence of the shape and material of the discharge chamber on the integral characteristics of the thruster was studied. The recommendations on the optimization of the thruster primary performance were developed based on the results of experimental studies.

A nonnegligible fraction of the charged particles in the Hall current thruster plume completes the electrical circuit through the conductive wall of a ground-based vacuum test facility. The resultant electrical circuit is different from the electrical circuit completed by the Hall current thruster in the onorbit environment. To understand the electrical circuit created in ground-based testing, this work examines the effect of an electrically biased metal plate, placed in the far-field plume of a Hall current thruster, on the plasma plume characteristics, the Hall current thruster thrust, and the electron termination pathways. An Aerojet Rocketdyne T-140 Hall current thruster is operated at 300 V and 10.3 A on xenon propellant. The operational neutral background pressure is 7.3×10−6 torr, corrected for xenon. Two aluminum plates, one representative of the radial wall of the vacuum chamber and one representative of the axial wall of the vacuum chamber, are placed 2.3 m radially outward from the thruster ce...

This paper introduces a concept of propellant-less microwave thruster for upper stage and space applications. The propellant-less microwave thruster works on classical electromagnetic theory, where the microwave radiations are converted to the electromagnetic pressure that acts as axial thrust. In microwave thrusters, microwaves are radiated to a sealed conical cavity which follows the contour of the thrusters to produce pressure. Normally, there are three types of methods by which thrust can be produced without utilizing any propellant. Cannae Drive, EmDrive, and the Wrap Drive are mostly used. The first two are based on the principle of microwave radiation, whereas the third method is just a concept in various science fiction films. The EmDrive can produce a small amount of thrust, therefore, its application is limited to upper stage thruster, orbit/trajectory correction, and deep-space exploration. This paper provides an insight on the feasibility of EmDrive for space application by comparing its performance with ion propulsion system. The proposed microwave propulsion system has low system mass (typically in Kgs), produces high thrust (in mN) and even has higher specific impulse (in years) which can outweigh the current ion propulsion system.

In order to develop multi-mode thrusters for all-electric propulsion spacecraft, a double-stage Hall thruster is designed and investigated experimentally. By applying two independent non-emissive electrodes and a double-peaked magnetic field in the double-stage Hall thruster, the propellant can be ionized in the first stage and then accelerated in the second stage. Experiment results revealed that double-stage discharge effectively improves thruster performance in high specific impulse working mode by reducing the electron current. A one-dimensional model is used to discuss the effects of discharge parameters on ionization distribution theoretically. As a conclusion, ionization distribution (in terms of position and length) in high specific impulse mode (mode I) is significantly different compared to high-thrust working mode (mode F). This double-stage Hall thruster provides a double-stage discharge in mode I to enhance the ionization in the channel and a single-stage discharge in mode F to reduce the ion wall-loss, achieving high efficiency in both working modes.

One of the most recently developed modes of propulsion is electric propulsion. The commonly used chemical propulsion systems have the advantage of a high Specific Impulse as compared to that of ion propulsion systems. However, owing to the efficacy of ion propulsion systems, it is considered the future of space exploration.Electro thermal thrusters produce thrust by using electrical fields to force hot plasma out of the nozzle with certain exit velocity. The plasma’s exit velocity and the system’s thrust capacity, as of now, are insufficient for space travel to be conducted within a reasonable time. I intend to study the possibility of improving the thruster’s performance by using an aerospike nozzle as an exit nozzle which meets the conditions required for the thruster to function appropriately. I shall be studying the plasma plume exit velocity variation with respect to the nozzles used. Also, a thermal analysis will be conducted in order to find the correct material for the nozzle.

The impact of facility conductivity on Hall effect thruster cathode coupling is experimentally investigated. The 3.4 kW Aerojet Rocketdyne T-140 Hall effect thruster operating at a discharge voltage of 300 V, a discharge current of 10.3 A, and an anode flow rate of 11.6 mg/s serves as a representative Hall effect thruster test bed. The nominal facility operating pressure during thruster operation is 7.3×10−6 Torr corrected for xenon. Two 0.91×0.91 m square aluminum plates are placed adjacent to, but electrically isolated from, the walls of the conductive vacuum chamber at two locations with respect to the center of the thruster exit plane: 4.3 m axially downstream along the thruster centerline, and 2.3 m radially outward centered on the exit plane. The plates and body of the Hall effect thruster are configured in three distinct electrical configurations with corresponding measurements: 1) electrically grounded with measurements of currents to ground, 2) electrically isolated with measurements of floati...

The physical process of a single-stage planar-pulsed-inductive accelerator is investigated. Measurements include the waveforms of circuit current, capacitor voltage, plasma radiation intensity, and temporal plasma structure photos captured by a high-speed camera. Experiments are conducted under static ambient fill condition using argon as propellant. Varied values of capacitor voltage and gas pressure are compared. Further discussions quantify the EM interaction between circuit and plasma, as well as their energy deposition and current sheet acceleration. Based on the results of experiments, physical mechanisms of the initial ionization phase and the following acceleration phase are analyzed theoretically.

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.

Electric propulsion is now a successful method for primary and secondary propulsion of deep space long-duration missions and for geosynchronous satellite attitude control. The Plasma Physics Laboratory of UnB has been developing a permanent magnet Hall thruster (PHALL) for the UNIESPACO program, part of the Brazilian space activities program (PNAE) since 2004. The idea of using an array of permanent magnets, instead of an electromagnet, to produce a radial magnetic field inside the plasma channel of the thruster is very significant. It allows the development of a Hall thruster with power consumption low enough to be used in small- and medium-size satellites. The PHALL project consists on plasma source design, construction and characterization of the Hall-type propulsion engine using several plasma diagnostics sensors. PHALL is based on a plasma source in which a Hall current is generated inside a cylindrical channel with an axial electric field produced by a ring anode and a radial magnetic field produced by permanent magnets. In this work, a brief description of the plasma propulsion engine, its diagnostics instrumentation and measured plasma parameters with a focus for possible applications of PHALL on orbit transfer maneuvering for future Brazilian geostationary satellite space missions is shown. More specifically, we will show plasma density and temperature space profiles inside and outside the thruster channel, ion temperature measurements based on Doppler broadening of spectral lines and ion energy measurements. Based on the measured plasma parameters we construct an aptitude figure of the PHALL. It contains the specific impulse, total thrust, propellant flow rate and power consumption necessary for orbit raising of satellites. Based on previous studies of geosynchronous satellite orbit positioning we perform numerical simulations of satellite orbit raising from an altitude of 700–36,000 km using a PHALL operating in the 100–500 mN thrust range. To perform these calculations integration techniques were used. The main simulation parameters were orbit raising time, fuel mass, total satellite mass, thrust and exhaust velocity. We conclude by comparing our results with results obtained with known space missions performed with Hall thrusters.

A lateral jet thruster operates at high temperature and high pressure, which causes ablation at the nozzle neck and malfunction of the nozzle system. The function of the solid propulsion system's nozzle is to provide thrust to the system by effectively utilising the high temperature gas that is generated in the combustion chamber. The requirements of the lateral propulsion system have to be taken into account in nozzle design, as well as the structural safety of the nozzle itself. An accurate stress analysis therefore needs to be conducted beforehand, simultaneously taking into account the thermal stress caused by the heat transfer between the high temperature gas and the wall, and the mechanical stress resulting from the pressure in the flow path. This study examined nozzle safety according to thruster operating time, where high temperature and high pressure combustion gas explodes to propel the prototype lateral thruster.

A new model of the plasma plume from Hall effect thrusters (HETs) has been developed for the purpose of more accurately predicting the interactions between future high-power thrusters and large high-voltage solar arrays, such as those being developed under NASA’s Game Changing Technology awards by ATK and Deployable Space Systems. The HET plume mainly consists of two types of ions. The first are the energetic main beam ions produced upstream of the thruster acceleration zone. These are the dominant ion species along the thrust axis. The other group of ions have lower kinetic energy and are generated downstream of the acceleration zone from neutral xenon gas atoms by charge exchange (CEX) with main beam ions and by electron impact ionization. The neutral gas is due to neutral propellant atoms leaving the thruster and the hollow cathode without being ionized, and, in the case of laboratory testing, background neutrals present in the vacuum chamber. The new model uses the 2-D axisymmetric Hall thruster code, Hall2De, to self-consistently calculate the three major components of the plume in the vicinity of the thruster, which are the neutral gas atoms, high-energy beam ions, and low-energy ions. From the boundary computational region in the near plume, the Hall2De results are propagated to distances of tens of meters using a continuum hydrodynamic algorithm. This approach offers important advantages with respect to prior models of Hall thruster plumes, such as the NASA Electric Propulsion Interactions Code (EPIC), which uses an analytical fit to laboratory data from a single thruster for the main beam velocity boundary conditions at the channel exit. EPIC assumes that the neutral gas density emanates uniformly and isotropically from the channel exit. Low-energy ions are generated only by CEX; low-energy ions generated by electron impact are not included. The results for the far-field plume of a conceptual high-power thruster (H6) show important differences between EPIC and simulations using the new plume model. High-energy ions undergo less expansion in the azimuthal direction than in EPIC. This can be attributed to magnetic focusing of the beam. We also observe that the peak in low-energy ion density at approximately 90° from the thrust axis predicted by EPIC is moved downstream to angles from 70° to 80° when the new plume model is employed.

A new type of pulsed plasma thruster (PPT) for small satellite propulsion is investigated, of which the most innovative aspect is the use of a non-volatile liquid propellant. The thruster is based on an open capillary design. The thruster achieved a thrust-to-power ratio above 45 μN/W, which constitutes a 5-fold improvement over the water-propelled pulsed plasma thruster, and which is also slightly above the performance of a similarly sized PPT with a solid propellant.

The cusped field thruster is an emerging kind of electric propulsion equipment employing a cusped magnetic field to achieve potentially long lifetime. As important factors contributing to the thruster performance, the optimal channel length and magnetic field design are still uncertain. In order to research how magnet length affects performance, a variable magnet length cusped field thruster (VML-CFT) was designed and tested in Harbin Institute of Technology (HIT). By using modular magnets, VML-CFT can vary the length of each cusp stage. The design of VML-CFT is demonstrated in this paper. Experimental results show that certain length increase of the middle stage can improve the thruster performance. By analyzing experimental results with different magnet lengths, the thruster internal physical process can be inferred.

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 Alfvenic flow transition ( $\beta _{k}$ ). There was no significant change observed to the applied magnetic field.

The scaling of magnetically shielded Hall thrusters to low power is investigated through the development and fabrication of a 4-cm Hall thruster. During initial testing, the magnetically shielded miniature Hall thruster was operated at 275 V discharge voltage and 325-W discharge power. Inspection of the channel walls after testing suggests that the outer discharge channel wall was successfully shielded from high-energy ion erosion while the inner channel wall showed evidence of weaker shielding, likely due to magnetic circuit saturation. Scanning planar probe measurements taken at two locations downstream of the thruster face provided ion current density profiles. The ion current calculated by integrating these data was 1.04 A with a plume divergence half-angle of 30°. Swept retarding potential analyzer measurements taken 80-cm axially downstream of the thruster measured the most probable ion voltage to be 252 V. The total thruster efficiency was calculated from probe measurements to be 43% (anode efficiency of 59%) corresponding to a thrust of 19 mN at a specific impulse of 1870 s. Discharge channel erosion rates were found to be approximately three orders of magnitude less than unshielded Hall thrusters, suggesting the potential for a significant increase in operational life.

To extend the lifetime of commercial and scientific satellites in low Earth orbit (LEO) and below (100–250 km of altitude) recent years showed an increased activity in the field of air-breathing electric propulsion as well as beamed-energy propulsion systems. However, preliminary studies showed that the propellant flow necessary for electrostatic propulsion at these altitudes exceeds the mass intake possible within reasonable limits, and that electrode erosion due to oxygen flow might limit the lifetime of eventual thruster systems. The pulsed plasma thruster (PPT), however, can be successfully operated with smaller mass intake and at relatively low power. This makes it an interesting candidate for air-breathing application in LEO and its feasibility is investigated within this paper. An analysis of such an air-breathing PPT system shows that for altitudes between 150 and 250 km, drag compensation is at least partially feasible assuming a thrust-to-power ratio of 30 mN/kW and a specific impulse of 5000 s. Further, to avoid electrode erosion, inductively heated electrothermal plasma generator technology is discussed to derive a possible propulsion system that can handle gaseous propellant without unfavorable side effects. Current technology can be used to create an estimated 4.4 mN of thrust per 1 mg/s of mass flow rate, which is sufficient to compensate the drag for small satellites in altitudes between 150 and 250 km.

Helicon plasma thruster is a novel electromagnetic thruster, which is based on the currentless double layer effect of the helicon plasma existing in the diverging magnetic field, accelerates the ions to form ion beam flow at a high speed, thus produces thrust. The key components of the HPT include a radio frequency antenna with specific structure and a magnetic field with certain type. This paper firstly summarizes the performance characteristics of the HPT, then gives the principle prototype design scheme for ground experiment of helicon plasma thruster concretely. The conclusion can provide the theoretical basis and reference for the further experimental study and the optimization design of the thruster.

The near-term potential for iodine propellant in Hall thrusters is explored. The merits of iodine with respect to other propellants are presented. Recent performance measurements are summarized, and new measurements taken with an 8 kW thruster are presented. Thruster discharge power exceeded 10 kW, and peak measured anode efficiency exceeded 65%. Spacecraft interactions issues are also addressed and relevant data taken with a 1 kW thruster are presented. Plume data showed lower divergence with iodine than with xenon, and that a plume shield could effectively attenuate the far field plume. Material samples placed in the plume showed a strong reaction with iron, but little reaction with typical spacecraft materials. System level benefits, including low storage pressure and extremely high density, are also discussed. All results so far indicate iodine is a viable propellant for electric rockets, and for some missions is superior to xenon.

The Hall effect thruster is an electric propulsion device for space applications that efficiently reduces the propellant mass requirements on a spacecraft in comparison with chemical rockets. To date, the Hall effect thruster technology relies on the thermionic cathodes that consume up to 10% of the total propellant used in the system to neutralize the ion beam of the thruster; however, such propellant usage does not contribute to thrust generation. An array of thin-film, carbon nanotube field emitters that emit electrons through field emission can potentially neutralize the ion beam without consuming any propellant. This paper examines the effects of 40 min exposure of carbon nanotube field emitter arrays to the plasma environment in the exit plane of a Hall effect thruster. The physical structures that enable field emission appear largely unaffected by placement in the plasma as well as operation in the plasma. This indicates that a refined design of this carbon nanotube field emitter array may potentially provide an alternative to the thermionic cathode used on contemporary Hall effect thrusters and verifies that no fundamental incompatibilities exist between these two technologies.

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.

In support of our planar pulsed inductive plasma thruster research, a fast electromagnetic inductive valve for a gas propellant injection system has been built and tested. A new and important design feature is the use of a conical diaphragm as the action part, which greatly contributes to the virtue of simplicity for adopting the resultant force of the diaphragm deformation as the closing force. An optical transmission technique is adopted to measure the opening and closing characters of the valve while the gas throughput is determined by measuring the pressure change per pulse in a test chamber with a capacitance manometer. The experimental results revealed that the delay time before the valve reaction is less than 40 μs, and the valve pulse width is no longer than 160 μs full width at half maximum. The valve delivers 0-2.5 mg of argon gas per pulse varied by adjusting the drive voltage and gas pressure.

This study adopts double permanent magnetic rings and four permanent magnetic rings to form two symmetrical magnetic peaks and two asymmetrical magnetic peaks in the channel of a Hall thruster, and uses a 2D-3V PIC-MCC model to analyze the influence of magnetic strength on the discharge characteristic and performance of Hall thrusters with an intermediate electrode and double-peaked magnetic field. As opposed to the two symmetrical magnetic peaks formed by double permanent magnetic rings, increasing the magnetic peak value deep within the channel can cause propellant ionization to occur; with the increase in the magnetic peak deep in the channel, the propellant utilization, thrust, and anode efficiency of the thruster are significantly improved. Double-peaked magnetic field can realize separate control of ionization and acceleration in a Hall thruster, and provide technical means for further improving thruster performance.

The increase in performance resulting from optimization of the magnetic field in a low-power magnetically shielded Hall thruster is investigated. The magnetic circuit of the MaSMi-60 Hall thruster was modified to improve the magnetic field topology while increasing the strength of the field across the discharge channel gap. Direct thrust measurements were then taken to assess the changes to thruster efficiency, thrust, and specific impulse. The MaSMi-60's total efficiency increased by nearly 30% as a result of the improved magnetic field, resulting in a peak value of 32.1% (38.6% anode efficiency). Peak thrust and total specific impulse values of 35.8 mN and 1,440 s (1,550 s anode specific impulse) were observed. To demonstrate the thruster's enabling capabilities when paired with a smallsat-class spacecraft, three example mission trajectories to 118401 LINEAR, an icy asteroid-belt comet, were calculated. For each trajectory, the MaSMi-60's experimentally demonstrated performance was used for the throttling table inputs. The trajectory solutions show a delivered mass fraction of between 35–49% for an initial spacecraft mass of up to 350 kg, a solar array power of up to 2.0 kW, and a total transfer time of ∼6.5 years.

The ion beam shepherd is an innovative contactless technique for space debris removal in which an impulse transfer thruster pushes the debris object through the action of a plasma plume and an impulse compensation thruster maintains formation flying. The optimal operational point of both thrusters strongly depends on their characteristics and on the physics of the plasma plume expansion into vacuum. With the use of dedicated thruster performance models, complemented with simplified plume expansion and plasma-debris interaction models, a system-level optimization study of the impulse transfer thruster alone and of the overall electric propulsion subsystem is presented for an ion beam shepherd mission example. An optimum design point is found for minimum overall power consumption in both cases. The research leading to the results of this paper was carried out within the LEOSWEEP project and received funding from the European Union Seventh Framework Programme (FP7/2007-2013) under grant agreement number 607457. Additional funding was received by Spain's Research and Development National Plan, grant ESP2013-41052

We have developed a three-dimensional particle model for a miniature microwave discharge ion thruster to elucidate the mechanism of ECR discharges confined in a small space. The model consists of a particle-in-cell simulation with a Monte Carlo collision algorithm (PIC-MCC) for the kinetics of charged particles, a finite-difference time-domain method for the electromagnetic fields of 4.2?GHz microwaves, and a finite element analysis for the magnetostatic fields of permanent magnets. The PIC-MCC results have shown that the electrons are well confined owing to the mirror magnetic fields and can be effectively heated in the ECR layer downstream of a ring-shaped antenna. The confinement results in the ring-shaped profiles of the plasma density along the antenna. The visual appearance of the plasma discharge of the thruster in operation was also ring-shaped. Moreover, the ions are expected to be accelerated effectively through the grid electrode without a large loss of ions toward side walls, that is, the plasma source developed here would be desirable in ion thrusters.

Research of magnetoplasmadynamic Pulsed Plasma Thrusters (PPTs), or iMPDs, at the Stuttgart Institute of Space Systems (IRS), led to the thruster design ADD SIMP-LEX. For optimization of the thruster's discharge behavior and of the plasma acceleration, the magnetic self-field is measured. Inductive miniaturized probes are introduced and used to acquire data at points of the volume inside of the thruster's discharge channel. The results are used for analysis to identify the extend of the discharge zone along the discharge channel of the PPT. Measured magnetic field signals along the centerline of the discharge channel are presented. The underlying distribution of the dynamic discharge current is deduced from Amperes circuital law and interpreted in a two-dimensional plot against camera images for the current-plasma-interaction, current motion and validity of today's discharge models. The evaluation of the calculated integral discharge current in a plane of the discharge channel against the recorded thruster discharge current is presented and discussed.

Currently, the propellant most often used in Hall effect thrusters is xenon and in recent years occasionally krypton. However, both noble gases are extremely scarce on Earth and missions with high propellant demand, e.g., high-power operation, are defied due to the limited annual production. Therefore, alternative abundant propellants like argon need to be considered. In this paper, the operation characteristics of the anode-layer Hall thruster (TAL) UT-58 with argon propellant are evaluated. Further, the guidelines of high-power operation with argon propellant have been investigated by discussing the discharge stability in each propellant using the same performance map. Finally, it is shown by these maps that optimum magnetic lines by magnetic shielding can be applied to the anode-layer Hall thruster with argon propellant. As a result, three regions were determined by comparing the oscillation amplitude using both xenon and argon, and it was discovered that the optimum operating points with argon propellant are in a region with low magnetic flux density.

Experimental results and a procedure for determining the total energy balance for single-stage thrusters with azimuth electron drift are presented. A single-stage thruster with an anode layer (TAL) with power of 200–800 W was used as an experimental model. Working medium was xenon. The near-anode potential drop was determined experimentally. The mean energy of beam ions was also determined by using a retarding field method and by integrating the retarding performances. The total ion beam current and power (useful power of TAL) are determined. The investigations are performed under minimal heat flux to an anode and for two geometries of magnetic system’s poles, which characterize different longitudinal gradients of magnetic field induction in a gas-discharge channel. It is revealed experimentally that thruster’s efficiency drops if the gradient of magnetic field induction in the channel increases. The results of investigation can be spread to more powerful thrusters.