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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.

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.

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)

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.

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.

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.

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 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.

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.

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

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.

The pulsed plasma thruster of the magnetoplasmadynamic type represents a promising and cheap solution to today’s propulsion challenges in space. The work presented in this paper covers new information on the electromagnetic conditions during thrust generation by revealing the temporal and spatial distributions of the current density vectors between the thruster electrodes. This is achieved by building and calibrating miniaturized induction probes. The local magnetic field components of all three spatial directions are combined to derive the current density vector. Quantitative determination of the temporal behavior of vectors at 27 points evenly distributed in three planes delivers a comprehensive picture of the position and development of the current in a pulsed plasma thruster. Maximum current densities of 80 kA/m2 and magnetic flux densities of up to 0.82 T have been observed close to the surface of the solid Teflon® propellant. It is concluded that no distinct movement of a current sheet was observed...

The role of the electrically conductive vacuum chamber wall in the completion of the discharge circuit of a Hall effect thruster is experimentally investigated. The Aerojet Rocketdyne T-140 laborat...

The helicon Hall thruster is a two-stage thruster that was developed to investigate whether a radio frequency ionization stage can improve the overall efficiency of a Hall thruster operating at high thrust and low specific impulse. This paper describes an experiment that measured the single-stage and two-stage performance of the helicon Hall thruster operating at 10–25 mg/s anode mass flow rates of xenon at 100–200 V discharge voltages, and also for 6 mg/s of argon at 300 V, and 2.6 mg/s of nitrogen at 200 V. The helicon Hall thruster performance during operation with argon and nitrogen is characterized by low beam divergence efficiency and low propellant utilization efficiency. During two-stage operation, the thrust of the helicon Hall thruster marginally increased with radio frequency power, but the propulsive efficiency and thrust-to-power both decreased with increasing radio frequency power. Probe diagnostics suggest that gains were realized by a slight increase in propellant efficiency, but that t...

This paper describes measurements of performance characteristics of micro-cathode arc thruster. In particular, influence of an applied magnetic field B on the impulse bit, specific impulse, and the mass consumption is investigated. The micro-cathode arc thruster was operated in a quasi-steady-state mode to ensure that a measurable thrust stand deflection was produced. It was found that the impulse bit increases from about 0.1 μN·s in the case without a magnetic field up to about 1.1 μN·s in the case of a magnetic field of about 0.3 T. A magnetic field leads to an increase of the mass consumption from about 10 μg/C at zero magnetic field up to about 50 μg/C at a magnetic field of about 0.3 T. Average ion velocity is measured to be about 18 km/s (at B=0 T) and about 31 km/s (at B=0.3 T). Maximum efficiency of the thruster is estimated to be about 7%.

The cusped field thruster uses several permanent ring magnets of alternating polarity to create a multistage cusp magnetic topology. It has demonstrated a long lifetime because of reduced plasma losses to the discharge channel surface. However, the high beam divergence problem has not been solved in this type of thruster. The purpose of this technical note is to describe a plume control method to reduce its plume angle. The results show that the magnetic field intensity in the plume region and the shape of the last separatrix have great influence on plume divergence.

A current-model Triple Langmuir probe was developed and used to measure electron temperature and density of the Pulsed Plasma Thruster plume. To decreasing the errors in measurement, Probe, collection circuit and glow cleaning devices were elaborately designed. A FIR digital Filter based on Matlab was designed and the software for date processing was developed by Labview. Measurements were taken at various position in the plume of a Pulsed Plasma Thruster operating at discharge energy of 6-24J. The results show the thruster plume has electron temperatures in the range between 0.6and 5.4eV, electron densities between 0.9×1019 and 4.1x1021m-3 for all discharge energy levels considered. Electron temperature and density decrease with increasing angle away from the centerline and with decreasing discharge energy.

The performance of a radio-frequency plasma thruster is evaluated using a displacement-type, inverted-pendulum thrust stand in the Large Vacuum Test Facility at the University of Michigan. A radio-frequency generator supplies up to 2000 W to the radio-frequency plasma thruster at a fixed frequency of 13.56 MHz. The matching network is placed inside the vacuum chamber at the thruster, and the radio-frequency power is measured at the matching network input port with a dual-directional coupler. A Faraday probe measures the current density in the far-field exhaust of the thruster, a retarding potential analyzer characterizes the ion voltage distribution, and a Langmuir probe measures the plasma potential. The radio-frequency plasma thruster is operated with argon propellant at mass flow rates of 2.4–7.6 mg/s, at corresponding corrected facility pressures of 1.5×10−6–4.6×10−6 torr. The maximum thrust observed is 10.8 mN, and the maximum specific impulse is 303 s. Thrust efficiency is below 1% at all conditio...

A thrust imparted by a permanent-magnets helicon plasma thruster is directly measured by using a pendulum thrust balance. The source consists of a 6.5-cm-inner diameter glass tube and a magnetic no...

Development of Hall thrusters for nano, small and low power satellites below 100 W is expected. In lowering Hall thruster power, the cylindrical type Hall thruster is more advantage than conventional coaxial-type Hall thrusters. In this study, a very low power cylindrical Hall thruster for nano-satellite “PROITERES-3” under development in Osaka Institute of Technology was designed, and the thruster performance was measured. As a result, a stable operation was achieved even with 10 W. The anode specific impulse and the anode thrust efficiency were 1570 s and 18.1%, respectively, with 66 W. Also, the discharge current oscillation was lower compared with that of SPT-type Hall thruster.

Performance measurements have been obtained of a novel propulsion concept called the Halo thruster under development within the University of Surrey. The Halo thruster, a type of cusped-field thruster with close similarity to the cylindrical Hall thruster, is motivated by the need for low-power and low-cost electric propulsion for the small satellite sector. Two versions of the device are investigated in this study: a design using permanent magnets at high magnetic-field strength and a design using electromagnets with moderate field strength. While operating at 200 W discharge power, which is of particular interest to power-limited small satellite platforms, the permanent-magnet design achieved a maximum thrust efficiency of 8% at a specific impulse of approximately 900 s using a krypton propellant. By comparison, the electromagnet design achieved a maximum thrust efficiency of 28% at a specific impulse of approximately 1500 s at 200 W using a xenon propellant. For higher levels of power (tested up to 800 W), the performance of the electromagnetic design saturated at approximately 25% thrust efficiency using krypton and 30% using xenon. The thrust efficiency of the permanent-magnet design appeared to increase monotonically up to 600 W reaching a maximum value of 14%.Read More: https://arc.aiaa.org/doi/abs/10.2514/1.B36091

NASA's Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) has been the subject of extensive technology maturation in preparation for flight system development. Part of the technology maturation effort included experimental evaluation of the TDU-1 thruster with conducting and dielectric front pole cover materials in two different electrical configurations. A graphite front magnetic pole cover thruster configuration with the thruster body electrically tied to cathode, and an alumina front pole cover thruster configuration with the thruster body floating were evaluated. Both configurations were also evaluated at different facility background pressure conditions to evaluate background pressure effects on thruster operation. Performance characterization tests found that higher thruster performance was attained with the graphite front pole cover configuration with the thruster electrically tied to cathode. A total thrust efficiency of 68% and a total specific impulse of 2,820 s was demonstrated at a discharge voltage of 600 V and a discharge power of 12.5 kW. Thruster stability regimes were characterized with respect to the thruster discharge current oscillations and with maps of the discharge current-voltage-magnetic field (IVB). Analysis of TDU-1 discharge current waveforms found that lower normalized discharge current peak-to-peak and root mean square magnitudes were attained when the thruster was electrically floated with alumina front pole covers. Background pressure effects characterization tests indicated that the thruster performance and stability were mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 1×10-5 Torr-Xe (for thruster flow rates of 20.5 mg/s). Power spectral density analysis of the discharge current waveforms showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant breathing mode frequency. Finally, IVB maps of the TDU-1 thruster indicated that the discharge current became more oscillatory with higher discharge current peak-to-peak and RMS values with increased facility background pressure at lower thruster mass flow rates; thruster operation at higher flow rates resulted in less change to the thruster's IVB characteristics with elevated background pressure.

In order to study the influence of electrode flare angle on the performance of Ablation Pulsed Plasma Thruster. Discharge character, plasma velocity and performance over different electrode flare angles of the Pulsed Plasma Thruster at 13.5J initial energy are measured experimentally, the effect of electrode flare angle on the electrode inductance gradient, equivalent circuit parameters and energy transfer efficiency are analyzed. It can be seen that the circuit parameters and inductance distribution are changed with electrode flare angle, the impulse bit, specific impulse, thruster efficiency and mean exhaust velocity increase non-linearly with flare angle increasing from 0 degree to 27 degree and the Pulsed Plasma Thruster gets the maximum thruster performance at 27 degree flare angle. It shows that with the increase of electrode flare angle the fraction of ablated mass accelerated magnetically and the impulse bit created by Lorentz force decrease are the reasons inducing the change in thruster performance.

The T-220HT is a 10-kW class Hall effect thruster developed as the primary propulsion system for satellites. In-channel electrodes and additional magnetic coils are added to study ion focusing to decrease energy losses from ion-wall neutralization and plume divergence in order to increase the thrust-to-power ratio. In this study, electrically-biased graphite electrode rings are embedded in the discharge channel walls to repel radial ions. The thruster is tested from 125-300 V at 9 A discharge, with the electrodes either floating, biased to 10 V or 30 V. The mass flow rate was varied from 9.8-10.4 mg/s to maintain constant current. Maximum chamber pressure was 1.5e-5 Torr-Xe. Performance measurements on xenon show a maximum increase in thrust-to-power ratio of 4.84 mN/kW, 15.3 mN thrust, 206 s Isp, and 8% anode efficiency. The plume ion current density, ion energy distribution function, and plasma potential is characterized and indicates a collimation of the ion beam and a increase in ion number density without an increase in propellant neutrals, which results in an increase in mass utilization. The different electrode currents and ion energy distribution functions at 10 V compared to 30 V electrodes leads to the idea of different modes of operation with different electrode biases.

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.

Nomenclature B, B = magnetic induction, T C = capacitance, F E = electric field, V=m E0 = initial energy, J e0 = specific energy, J=kg I = current, A Ibit = impulse bit, N s Isp = specific impulse, s j = current density, A=m _ L = dynamic impedance, H=s LC = coil inductance, H L0 = initial inductance, H L0 = inductance per unit length, H=m L = inductance ratio M = mutual inductance, H; molecular mass, kg mbit = mass bit, kg=pulse mc = critical mass, kg Q1 = first ionization potential, J r = radial coordinate, m Re = external resistance, Rp = plasma resistance, t = time, s ue = exhaust velocity, m=s v = velocity, m=s Vp = plasma voltage, V V0 = initial charge voltage, V z = axial coordinate, m z0 = electromagnetic stroke length, m = dynamic impedance parameter L = inductance change, H = resistivity, m; efficiency, % = nonpropulsive energy fraction A = linear mass density, kg=m = magnetic flux, W 1, 2 = critical resistance ratios

High thrust-to-power ratio on Hall-effect thrusters allows for more efficient thrust generation and better overall performance. It allows Hall-effect thrusters to perform more thrust-intensive maneuvers such as orbit raising, thus saving the mass of a second engine purely for those maneuvers. This paper presents the initial performance measurementsofaHall-effectthrusterusingin-channelelectrodestocollimateionsinthedischargechannelinorder to increase thrust-to-power. The thruster is tested in the range of 100–300 V discharge voltage, at 9 and 20 A discharge current, and with 0–30 V electrode potential on krypton propellant. Xenon propellant is used briefly. A null-type, inverted pendulum thrust stand is used to acquire thrust measurements. Improvements of 3 mN=kW, 200sspecificimpulse,and4%anodeefficiencyareobservedwithenergizedelectrode,supportingtheioncollimation hypothesis. Improvements exist primarily at low voltages (

The paper designs a pulsed plasma thruster system based on water propellant and studies energetic threshold of its main working characteristics to solve the problems of low performance and high pollution of Teflon pulsed plasma thrusters.Based on an analysis of the forming condition and cause of electric discharge types,working energetic threshold is studied experimentally by the discharge voltage and current measurement.The study shows that the initial plasma that can provide trigger probability for steady operation and energy storage capacity are obtained only by enough electric-field intensity and trigger current,and the normal working energy range of the water propellant pulsed plasma thruster is decided by stored energies but not related to capacity values.The study provides a possibility for developing low power water propellant pulsed plasma thrusters of free pollution,high performance and wide applications.