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An axial profile of the quasineutral plasma jet exhaust from a Hall thruster is studied. The plasma jet expansion is modeled using the sourceless steady-state hydrodynamic equations. It is considered that the plasma jet has a conical shape with a half angle of about 40°. The magnetic field surrounding the Hall thruster exit is included in the calculation. It is found that a magnetic field may significantly affect the axial profile of the plasma potential. For instance, in the case of zero magnetic field, the plasma potential is about −10 V at 1 m from the thruster exit, while in the case of a 0.1 T magnetic field, the plasma has a potential of about +25 V. Results predicted by the model are found to be in good agreement with experimental data for three different Hall thruster designs.

We propose a new method for the analysis of the efficiency of ion generation and acceleration, which can be used to determine the ion current in a plasma accelerator (thruster) with closed electron drift. The method is based on the measurement of the energy and angular distributions of ions in the plasma jet, the discharge power, and the jet thrust. The laws of variation of the plasma thruster efficiency as a function of the discharge voltage in the interval of 300–1000 V have been studied.

A numerical model of the ion beam in field emission electric propulsion thrusters is presented. The objective of the study is to develop a computational tool capable of predicting the effects of thruster configuration parameters (geometry, voltage distribution between the electrodes) on thruster performance (beam divergence, ion exhaust velocity). A two-dimensional, steady model has been adopted. A simple emission model has been assumed, neglecting the detailed, microscopic description of the physical emission mechanism. The electric field is evaluated by a finite element scheme; boundary conditions are provided by an auxiliary technique based on the placement of fictitious charges in proximity to the electrode surfaces. The beam is simulated through a discrete number of charged particles; the effects of the associated space charge density distribution on the electric field is considered. The results of the simulations of typical thruster configurations are presented and discussed. The need to take space charge effects into account to attain a sufficiently accurate solution is indicated: the beam charge density modifies the electric potential distribution generated by the electrodes voltage difference, causing a decrease in the local electric field at emitter tip and an increase in beam divergence. Neutralization at a finite distance must be taken into account to fully appreciate the effects of different voltage distributions between the electrodes: a higher accelerator electrode voltage determines a lower ion exhaust velocity and a larger beam divergence. The capability of an additional, decelerating neutral electrode to focus the ion beam is also investigated, showing that only little improvement in beam containment is attained.

The Hall current plasma thruster accelerates a plasma jet by an axial electric field and an applied radial magnetic field in an annular ceramic channel. A relatively large current density (>0.1 A/cm2) can be obtained as the acceleration mechanism is not limited by space charge effects. Such a device can be used as a small rocket engine on board spacecraft with the advantage of a large jet velocity compared to conventional rocket engines (10000–30000 m/s versus 2000–4800 m/s). An experimental Hall thruster was constructed and operated in a broad range of operating conditions and under various configuration variations. Electrical, magnetic and plasma diagnostics, and as well accurate thrust and gas flow rate measurements, have been used to investigate the dependence of thruster behavior on the applied voltage, gas flow rate, magnetic field, channel geometry and wall material. The studies conducted so far have demonstrated a significant effect of channel material on thruster electrical characteristics and th...

The oscillations from 0-100 kHz in a 6-kW magnetically shielded thruster are experimen- tally characterized. Changes in plasma parameters that result from the magnetic shielding of Hall thrusters have the potential to significantly alter thruster transients. A detailed investigation of the resulting oscillations is necessary both for the purpose of determin- ing the underlying physical processes governing time-dependent behavior in magnetically shielded thrusters as well as for improving thruster models. In this investigation, a high speed camera and a translating ion saturation probe are employed to examine the spatial extent and nature of oscillations from 0-100 kHz in the H6MS thruster. Two modes are identified at 8 kHz and 75-90 kHz. The low frequency mode is azimuthally uniform across the thruster face while the high frequency oscillation is concentrated close to the thruster centerline with an m = 1 azimuthal dependence. These experimental results are discussed in the context of wave theory as well as published observations from an unshielded variant of the H6MS thruster.

The thruster ADD SIMP-LEX (ADvanceD Stuttgart Impulsing Magneto-Plasmadynamic Thruster for Lunar EXploration) has been developed at the Institute of Space Systems (IRS) in Stuttgart, to provide secondary and primary propulsion capabilities to small satellites. Extensive optimization led to a mature design of world leading thrust efficiency around 30 % with russian APPT devices. The thruster was chosen for development into a stand-alone system in 2010. The system is capable of operating under satellite-like vacuum conditions in a test chamber, foregoing the usual outside laboratory equipment. The only connections required for system operation are a serial command interface, thermal housekeeping and a 28 V power line. Telemetry of the system for thruster health monitoring includes measurement of the pulsed current signal at up to four thruster capacitors. Requirements of the thruster and operation conditions ruled out any available COTS solutions for this task. A custom designed current measurement system was developed at the ETH Zurich and successfully integrated into the ADD SIMP-LEX design at IRS. The measurements revealed the current signal for every capacitor terminal and allowed for reconstruction of the combined flow. This data showed a current ringing effect between capacitors, which is undesirable for efficient thruster operation. The capacitor setup was changed and impact on the ringing was analyzed. The results showed reduction of the effect. Although complete elimination was not achieved, the data suggest careful consideration of the power train configuration can achieve satisfying mitigation levels. The sensor data quality was significantly improved by overcoming electromagnetic compatibility issues through introduction of a fiber-optical (LWL) transfer interface.

A unique configuration of the magnetic field and channel geometry near the wall of Hall thrusters, called magnetic shielding, has recently demonstrated the ability to significantly reduce the erosion of the boron nitride (BN) walls and extend the life of Hall thrusters by orders of magnitude. The ability of magnetic shielding to minimize interactions between the plasma and the discharge chamber walls in regions where erosion typically occurs has for the first time enabled the replacement of insulating walls with conducting materials without loss in Hall thruster performance. It is important to note that this is not a thruster with anode layer (TAL) where the walls are at or near cathode potential, but is a Hall thruster configuration where the walls are near the anode potential. The BN rings in the 6-kW H6 Hall thruster were replaced with graphite that self-biased to near the anode potential during operation. The thruster efficiency remained over 60% (within 2% of the baseline BN configuration) with a small decrease in thrust and increase in Isp typical of magnetically shielded Hall thrusters. The graphite wall temperatures decreased significantly compared with both shielded and unshielded BN configurations, leading to the potential for higher power operation. Eliminating ceramic walls makes it simpler and less expensive to fabricate a thruster to survive launch loads, and the graphite discharge chamber radiates more efficiently, which increases the power capability of the thruster compared with conventional Hall thruster designs.

Summary form only given. In this study, a SPPT (Solid Propellant Pulsed Plasma Thruster) driven by a MPC (Magnetic Pulse Compression) pulse generator is presented. Electromagnetic SPPT is well-known and is one of the most simple structures for electric propulsion: it is very lightweight and reliable and its thrust force can be easily controlled by the pulse discharge characteristics. In this type of system propellants are supplied from ablation of the solid propellant by surface discharge; therefore, it is difficult to regulate the propellant supply and ensure efficient use. This causes low performance of PPT. To achieve efficient utilization of the propellant, a MPC pulse generator was used as the power source of the SPPT. The magnetic switch which was used in the MPC has a magnetic saturation characteristic. When the magnetic switch is unsaturated, discharge current is restricted by the high inductance of the magnetic switch, and propellants are fed and excited by surface discharge. Later, when the magnetic switch is saturated, high current flows through the plasma and it is accelerated by the Lorentz force. Here, a high di/dt converted the mechanism to acceleration, quickly. It was contributed to reduce the Late-Time Vaporization by low rising temperature. Also, the main discharge was formed including pre-spread plasma and it accelerates the plasma efficiently. The pulse discharge of the MPC also composes the igniter which does not require an external power supply. A co-axial electrode type MPC-SPPT was designed and operated to demonstrate the system. The igniter was put as the center electrode and driven with a MPC pulse discharge. The 18 J MPC with a thruster body has a 41.6 nF energy storage capacitor and less than 80 nH of parasitic inductance. The peak discharge current is 15 kA, and the thrust performances were evaluated by a thrust target. From experimental results the advantages and capabilities of a MPC-SPPT system are discussed.

The Flux-Coil-Generated, Field-Reversed Configuration Experiment (FCX) at Tri Alpha Energy, Inc. (TAE) requires an initial source of plasma, from which the FRC is formed. This paper details the design, simulation, and performance measured for a new set of PITs developed for the FCX. The simulation solves coupled differential equations for the circuit, plasma parameters, including formation and acceleration. The PIT construction involves: 5 gas valves, a 2-stage Marx generator capacitor bank, a low-inductance parallel-plate transmission line and a H.V. feedthrough. The system delivers a 2.5 us, 1/4 period pulse and 1 Ohm impedance. Measurements at the surface of the PIT indicate: n = 1014-1015 cm-3, Ti = 30 eV, Vi = 100 eV. Measurements 1-m downstream indicate: n = 1012-1014 cm-3, temperature, Ti = 10 eV, and total plasma energy content, Wp = 30 J, which are adequate for the FCX.

The radio frequency ion thruster is an impulse generator to be used in space missions. It is a plasma based generator which utilizes electrostatic field between grids to accelerate the ionized gas out of the thruster to generate thrust. An analytical model is used to optimize the geometry of the discharge chamber which contains the inductively coupled plasma. The model is implemented and verified with existing thrusters and the results of the optimization process are presented. The results of our optimization routine validate the recent trend in radio frequency thruster design, which promotes hemispherical discharge chambers.

Summary form only given. The Particle-in-Cell (PIC) method was used to study various plasma thruster concepts: Stationary Plasma Thruster (SPT), High Efficiency Multistage Plasma Thruster (HEMP-T) and the cylindrical Hall thruster (CHT). In all cases PIC proved itself as a powerful tool, delivering important insight into the basic physics of the different thruster concepts. Two dimensional Particle-in-Cell code with Monte Carlo Collisions (PIC-MCC) was applied to simulate the operation of SPT and HEMP thrusters, in particular the plasma properties in the discharge chamber due to the different magnetic field configurations. Special attention was paid to the simulation of plasma particle fluxes on the thrusters' channel surfaces. The model resolves 2 spatial (rz) and 3 velocity components (2d3v). Using ion flux distributions calculated with PIC-MCC, the wall erosion of both thruster types was studied with the binary collision approximation (BCA) based Monte-Carlo code SDTrimSP1. In order to resolve the electron cross-field transport selfconsistently, accounting for plasma azimuthal dynamics, a new, fully three-dimensional PIC-MCC model was developed2. The model utilizes an equidistant Cartesian grid which explicitly assures momentum conservation and zero self forces. The simulation treats electrons, Xe+ ions and neutral Xenon atoms. All relevant collisional processes are included in the model. The dynamics of the background neutral gas is self-consistently resolved with direct simulation Monte Carlo. The model was applied to simulate the operation of a 100 W CHT thruster3. In the simulation a spoke rotating with the velocity of about 2 km/s was observed in the thruster channel, which agrees very well with the experiment4. The initial position of the spoke was found to be strongly correlated with the cathode placement. The simulation has shown that the rotating spoke is associated with the strong azimuthal depletion of the neutral gas, which indicates the importance of the ionization for this instability.

Among the various types of plasma thrusters developed over the last few decades for in-space propulsion applications of satellites and spacecrafts, Hall effect thrusters and ion engines are the most studied ones. The work at Bogazici University is concentrated on radio frequency ion thrusters. In this paper the design and manufacturing methods of the BURFIT-80 (Bogazici University Radio Frequency Ion Thruster-80) are explained. The BURFIT-80 is an experimental ion thruster and will be tested in the vacuum chamber at Bogazici University. The Ion Thruster has a cylindrical discharge chamber having an inner diameter and outer of diameter of 80 mm and 88 mm respectively, with a length of 72 mm. A double grid system with 91 holes is implemented. Preliminary calculations suggest that the thruster will generate thrust in the range of 2 to 6 mN. A discussion on material selection and manufacturing methods is presented for the production of the prototype.

A new design for a permanent magnet helicon thruster is presented. Its small plasma volume (∼10 cm3) and low power requirements (

Summary form only given. Electric propulsion is now a successful method for primary propulsion of deep space long duration missions and for geosynchronous satellite attitude control. The Closed Drift Thruster, also called Hall Thruster or SPT (Stationary Plasma Thruster) was primarily conceived in USSR (the ancient Soviet Union) and since then, it has been developed by space agencies, space research institutes and industries in several countries such as France, USA, Israel, Russian Federation and Brazil. In this work, we present the main features of the Permanent Magnet Hall Thruster (PMHT) developed at the Plasma Laboratory of the University of Brasilia. This project is supported by the Brazilian Space Agency program for universities named UNIESPACO. 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 description of a new vacuum chamber built to test the second prototype of the PMHT (PHALL II) is given. PHALL II has an aluminum plasma chamber and is smaller with 15 cm diameter and will contain rare earth magnets. We show the plasma density and temperature space profiles inside and outside the thruster channel. Ion temperature measurements based on the Doppler broadening of spectral lines and ion energy measurements are also shown. Based on the measured plasma parameters, we construct an aptitude figure of the PMHT. It contains the specific impulse, total thrust, propellant flow rate and power consumption necessary for satellites orbit raising and attitude control. Based on previous studies of geosynchronous satellite orbit positioning, we perform numerical simulations of satellite orbit raising from an altitude of 700 km to 36000 km using a PMHT operating in the 100mN-500 mN thrust range. In order 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 comparing our results with results obtained with known space missions performed with Hall Thrusters.

We describe the micro-thrust automatic measurement system of stationary plasma thruster in the article, The system can carry a thruster and power supply less than 3kg weight. Thrust capacity is from 1 to 100 mN,and thrust expend uncertainty is 0.2 mN（k=2）.The system is based on the torsion balance principle. The thrust of the thruster makes torque to the pivot. This torque is balanced by the electromagnetic compensate torque generated by the measurement system. The rotation angle of this system is measured by the position sensor, the electromagnetic interference single comes from the speed sensor, the electromagnetic compensate torque is from the torque sensor. This is a closed circuit and automatic measurement system. The additional forces generated by the weight, the air supply system and the wires of the thruster have been separated from the micro-thrust in this system. The influence of deadweight, air supply system and power supply system of thruster on exact measurement of micro-thrust have been eliminated, so micro-thrust of thruster can be measured precisely and automatically.

This chapter considers specifics of plasma application in space propulsion. Space propulsion is required for satellite motion in the outer space. The plasma physics and engineering of thrusters mainly based on electromagnetic plasma acceleration are described. Hall thruster, pulsed plasma thrusters, microthruster are some examples of plasma thrusters described in this Chapter. The basic physics of electron transport across the magnetic field, which is one of the most important phenomena in Hall thrusters, is considered. In the framework of the pulsed plasma thruster this Chapter covers the ablation phenomena in the presence of dense plasma.

Theoretical studies indicate that a variety of microscopic and macroscopic plasma instabilities can significantly affect the performance of the magnetoplasmadynamic (MPD) thruster, yet experimental evidence for their existence is extremely limited. The objective of this research was to provide experimental evidence for the existence of microinstabilities in a 10-kW-class, self-field MPD thruster. The lower- bound of the range of thruster operating currents investigated in this study was set by the requirement that the electromagnetic thrust component be much greater than the electrothermal thrust component. The upper-bound of the operating current was set by the initial observation of high-frequency thruster voltage oscillations, which is usually associated with the onset of excessive thruster erosion. Experimental evidence for two microinstabilities was obtained in the near-field plume of the thruster: 1) the generalized lower hybrid drift instability and 2) the electron cyclotron drift instability.

Several magnet assemblies relevant to ion thrusters were investigated using a numerical code to calculate the primary electron mean containment time. An analytical model was used to determine, consistently with the plasma stability condition, the ion thruster performance parameters, namely, the plasma ion energy cost, the plasma density, the plasma potential, and the doubly charged ion production rate as a function of the propellant mass flow rate. The numerical code was checked by reproducing the experimental performance parameters obtained from a 7-m checkerboard ion thruster. Using this approach, performance curves were obtained for a 5-cm ion thruster devised to produce a 1-mN thrust with low power and propellant consumption.

Results are presented of measurements with double and single electric probes in the beam of a UK-10 ion thruster, using xenon as the propellant gas. The probe material consisted of tungsten and secondary emission was of negligible importance. As a consequence of the large directed velocity of the ions, the ion current to a probe was essentially independent of probe voltage. Since the debye length was small compared with the probe radius, no ions or electrons were collected on the downstream part of a probe. The thruster was operated at a thrust level of 18 ± 0.3 mN. Detailed results are given for ion flux profiles and ion density in the beam, as well as for floating potential profiles and degree of nonneutralization. Ion densities reach a maximum of 1.8 ± 0.2 x 10'° cm~3 on the beam centerline at an axial distance of 15 ± 0.3 cm from the acceleration grid. The floating potential at that location is 8.5 ± 0.2 V with respect to the neutralizer, while the degree of nonneutralization there is on the order of 10~ 4. Characteristic electron energies vary from 0.5 to 2.7 eV. Evidence is obtained that the electron gas has a non-Maxwellian speed distribution function.

Plasma potentials and electron temperatures were deduced from emissive and cold floating probe measurements in a 2 kW Hall thruster, operated in the discharge voltage range of 200–400 V. An almost linear dependence of the electron temperature on the plasma potential was observed in the acceleration region of the thruster both inside and outside the thruster. This result calls into question whether secondary electron emission from the ceramic channel walls plays a significant role in electron energy balance. The proportionality factor between the axial electron temperature gradient and the electric field is also significantly smaller than might be expected by fluid models.

We demonstrate a solid fuel electrodeless ion thruster that uses a magnetic nozzle to collimate and accelerate copper ions produced by a high power impulse magnetron sputtering discharge (HiPIMS). The discharge is initiated using argon gas but in a practical device the consumption of argon could be minimised by exploiting the self-sputtering of copper. The ion fluence produced by the HiPIMS discharge was measured with a retarding field energy analyzer (RFEA) as a function of the magnetic field strength of the nozzle. The ion fraction of the copper was determined from the deposition rate of copper as a function of substrate bias and was found to exceed 87%. The ion fluence and ion energy increased in proportion with the magnetic field of the nozzle and the energy of the ions was found to follow a Maxwell-Boltzmann distribution with a directed velocity. The effectiveness of the magnetic nozzle in converting the randomized thermal motion of the ions into a jet was demonstrated from the energy distribution of the ions. The maximum ion exhaust velocity of at least 15.1 km/s, equivalent to a specific impulse of 1543 s was measured which is comparable to existing Hall thrusters and exceeds that of Teflon pulsed plasma thrusters.

The multi-cusped field thruster is a kind of electrostatic thruster used for spacecraft propulsion. The thruster typically operates in both high and low current modes with a visual change in the thruster plume. A Faraday probe and a retarding potential analyzer were employed to discover the structures of the hollow plume in these two modes. The results show that in the small angle region (from 0° to 15°), the ion energy distribution function in the high current mode contains two peaks. While in the low current mode, only the low-energy peak is remained. Besides, during the mode transition between these two modes, a reverse variation trend of ion energy distribution occurs in the middle angle region (from 15° to 40°), which indicates that there are two kinds of acceleration routes for the ions generated in the discharge channel. It is suggested that the change of ion current in these two routes could lead to the transition of the hollow plume between different operating modes.

The performance characteristics of a Multi-cusped Field Thruster depending on the magnetic field strength in the discharge channel were investigated. Four thrusters with different outer diameters of the magnet rings were designed to change the magnetic field strength in the discharge channel. It is found that increasing the magnetic field strength could restrain the radial cross-field electron current and decrease the radial width of main ionization region, which gives rise to the reduction of propellant utilization and thruster performance. The test results in different anode voltage conditions indicate that both the thrust and anode efficiency are higher for the weaker magnetic field in the discharge channel.

The azimuthal plasma current in a magnetic nozzle of a radiofrequency plasma thruster is experimentally identified by measuring the plasma-induced magnetic field. The axial plasma momentum increases over about 20 cm downstream of the thruster exit due to the Lorentz force arising from the azimuthal current. The measured current shows that the azimuthal current is given by the sum of the electron diamagnetic drift and drift currents, where the latter component decreases with an increase in the magnetic field strength; hence the azimuthal current approaches the electron diamagnetic drift one for the strong magnetic field. The Lorentz force calculated from the measured azimuthal plasma current and the radial magnetic field is smaller than the directly measured force exerted to the magnetic field, which indicates the existence of a non-negligible Lorentz force in the source tube.

Analysis is made for the possibility of erosion pulsed plasma thruster (PPT) application while solving the task of maintaining the point sustaining a long-operating geostationary artificial Earth satellite. The concept of an erosion PPT with a pulse energy of 200–300 J, designed for holding the attitude of a satellite of 500 kg mass over 10 years is presented. The thruster, with a lifetime of 2–3 × 10 7 pulses, produces a total pulse of 2.5 × 10 5 Ns, consuming up to 13 kg of propellant (Teflon). Estimations have shown that the thruster total mass will not exceed 50–60 kg for the flying PPT. Even in the case of a two-fold redundancy the thruster set mass should comprise no more than 0.20-0.25 of the satellite mass. The prospect of a rail design for the PPT discharge chamber with lateral propellant feeding is experimentally examined.

In order to minimize the leakage fields of a MHD thruster, we study a modular thruster composed of superconducting windings (SC) based on sheets to form "current boxes" or made of massive conductors, around an annular active channel. First, a 2D complex field formulation is used when SC are massive wires. Next, a bi-dimensional and a tri-dimensional electromagnetic modelization of the "current box" problem are developed using a direct method to calculate the field at every point. >

The multi-cusped field thruster is a novel concept in electric propulsion, which employs alternating permanent polarity magnets to create a periodic magnetic field. As an important factor for the optimal design, the magnet length of a low-power thruster is designed to be changeable. The effects of the ultimate stage length (L u) and the middle stage length (L m) on the performance characteristics are studied in a series of experiments. The results show that increasing L u could prolong the axial motion range of the electrons and promote the ionization process. In addition, it can enlarge the relative distance between the ionization and the acceleration regions. Both of these aspects can help improve the thrust and the anode efficiency. With regard to he reverse trend, wefound that the longer L m leads to a reduction of the thruster's performance for the enhanced ion loss to the wall. Overall this study can provide some optimal design ideas on the magnet length to improve the total performance of the multi-cusped field thruster.

Operation of a 5 kW-class Hall current Thruster for various voltages from 400 V to 800 V and a xenon mass flow rate of 6 mg s−1 have been studied with a quasi-neutral hybrid model. In this model, anomalous electron transport is fitted from ion mean velocity measurements, and energy losses due to electron–wall interactions are used as a tuned parameter to match expected electron temperature strength for same class of thruster. Doubly charged ions production has been taken into account and detailed collisions between heavy species included. As the electron temperature increases, the main channel of Xe2+ ion production becomes stepwise ionization of Xe+ ions. For an applied voltage of 800 V, the mass utilization efficiency is in the range of 0.8–1.1, and the current fraction of doubly charged ions varies between 0.1 and 0.2. Results show that the region of ion production of each species is located at the same place inside the thruster channel. Because collision processes mean free path is larger than the acceleration region, each type of ions experiences same potential drop, and ion energy distributions of singly and doubly charged are very similar.

The performance characteristics of a cylindrical Hall thruster depending on the magnetic field near the anode were investigated. The magnetic shielding rings were designed to adjust the magnetic field near the anode in different levels. The experiment results show that decreasing the magnetic field near the anode contributes to the enhancement of propellant utilization and the narrowing of ion energy distribution. It is suggested that the ionization region extends towards the anode, meanwhile, the angular distribution of ion beam is narrower, which could be attributed to the growing azimuthal current. As a result, the thrust and efficiency are enhanced significantly. This work can provide some optimal design ideas of the magnetic field to improve the performance of the thruster.

A pulsed axial magnetic field up to ∼2.8 kG is applied to a 26-mm-inner-diameter helicon plasma thruster immersed in a vacuum chamber, and the thrust is measured using a pendulum target. The pendulum is located 30-cm-downstream of the thruster, and the thruster rf power and argon flow rate are fixed at 1 kW and 70 sccm (which gives a chamber pressure of 0.7 mTorr). The imparted thrust increases as the applied magnetic field is increased and saturates at a maximum value of ∼9.5 mN for magnetic field above ∼2 kG. At the maximum magnetic field, it is demonstrated that the normalized plasma density, and the ion flow energy in the magnetic nozzle, agree within ∼50% and of 10%, respectively, with a one-dimensional model that ignores radial losses from the nozzle. This magnetic nozzle model is combined with a simple global model of the thruster source that incorporates an artificially controlled factor α, to account for radial plasma losses to the walls, where α = 0 and 1 correspond to zero losses and no magnetic field, respectively. Comparison between the experiments and the model implies that the radial losses in the thruster source are experimentally reduced by the applied magnetic field to about 10% of that obtained from the no magnetic field model.

資料番号: SA1160299000

We have investigated the plasma production in a micro RF ion thruster, where the plasma source is 5.0 mm in radius and 10 mm in length. To find the optimum condition to generate the RF plasma in the wide range of frequency f = 1-1000 MHz and pressure p = 0.1-10.0 Pa, we employ the equivalent circuit model, where the global model is also incorporated to obtain the plasma parameters self-consistently. The numerical results have indicated that capacitive coupling dominated over inductive coupling and the sheath resistance has a significant influence on the power coupling efficiency at lower frequency. The power coupling efficiency could be more than 80% at higher frequencies (> 30 MHz).

An innovative facility for testing high-power, pulsed plasmoid thrusters has been constructed to develop the electrodeless Lorentz force (ELF) thruster concept. It is equipped with a suite of diagnostics optimized to study the physical processes taking place within ELF and evaluate its propulsive utility including magnetic field, neutral gas, and plasma flux diagnostics, a method to determine energy flow into the plasma from the pulsed power systems, and a new type of ballistic pendulum, which enables thrust to be measured without the need for installing the entire propulsion system on a thrust stand. Variable magnetic fields allow controlled studies of plume expansion in a small-scale experiment and dielectric chamber walls reduce electromagnetic influences on plasma behavior and thruster operation. The unique capabilities of this facility enable novel concept development to take place at greatly reduced cost and increased accessibility compared to testing at large user-facilities.

A microwave discharge ion thruster of 5cm size was fabricated and its next concept improved in plasma ignition and sustention was reported. As we have already developed 10cm and 2cm size microwave discharge type ion thrusters, the development of 5cm one on the basis of the previous ones is carried out. It is sometimes difficult to generate plasma in it. Several magnetic configurations in the discharge chamber were calculated for the development of 5cm thruster with variable magnetic field.

Development of microwave discharge ion thrusters of 3cm size thruster was conducted to confirm the design approach to small scale microwave discharge ion thrusters, and concept of 5cm size thruster improved in plasma ignition and sustention was reported. As we have already developed 10cm and 2cm size microwave discharge type ion thrusters, the development of 3cm one on the basis of the previous ones is carried out. It is sometimes difficult to generate plasma in the 5cm size microwave discharge ion thruster. Several magnetic configurations in the discharge chamber were calculated for the development of 5cm thruster with variable magnetic field.

Plasma behavior analysis of a rectangular laser-assisted pulsed plasma thruster (LAPPT) was conducted. Laser-ablation plasma in the thruster was induced through laser beam irradiation onto a solid target and accelerated by electrical means instead of direct acceleration only by using a laser beam. Plasma behaviors between planar electrodes were observed with time-resolved ICCD camera imaging. Moreover, time-of-flight (TOF) measurement of ions in the exhaust plasma plume was also conducted. As the results from the ICCD imaging, typical speed of plasmas for charge voltage of 2000 V was 30 km/s. On the other hand, from the TOF measurement, it was over 100 km/s. Together with the discussion of this discrepancy, plasma acceleration mechanism was also discussed.

Operational characteristics of two separate inductive thrusters with conical theta pinch coils of different cone angles are explored through thrust stand measurements and time- integrated, unfiltered photography. Trends in impulse bit measurements indicate that, in the present experimental configuration, the thruster with the inductive coil possessing a smaller cone angle produced larger values of thrust, in apparent contradiction to results of a previous thruster acceleration model. Areas of greater light intensity in photographs of thruster operation are assumed to qualitatively represent locations of increased current density. Light intensity is generally greater in images of the thruster with the smaller cone angle when compared to those of the thruster with the larger half cone angle for the same operating conditions. The intensity generally decreases in both thrusters for decreasing mass flow rate and capacitor voltage. The location of brightest light intensity shifts upstream for decreasing mass flow rate of propellant and downstream for decreasing applied voltage. Recognizing that there typically exists an optimum ratio of applied electric field to gas pressure with respect to breakdown efficiency, this result may indicate that the optimum ratio was not achieved uniformly over the coil face, leading to non-uniform and incomplete current sheet formation in violation of the model assumption of immediate formation where all the injected propellant is contained in a magnetically-impermeable current sheet.

Drag-free control systems have been studied to detect gravitational waves by measuring distortions in the fabric of space-time. The purpose of drag-free control is to counteract nongravitational forces such as residual aerodynamic drag or solar radiation pressure. The objective of this study is to use miniature microwave discharge ion thruster, 1, for this propulsion system. In this experiment, the thrust dynamic range was between 7-100 %, and the thrust noise was less than 0.02 N/Hz 1/2 in the frequency range of 0.1-1.0 Hz. Although the design of 1 thruster has not yet been optimized, we can expect that the 1 thruster will be feasible for DPF thruster after optimization.

As NASA considers future exploration missions, high-power solar-electric propulsion (SEP) plays a prominent role in achieving many mission goals. Studies of high-power SEP systems (i.e. tens to hundreds of kilowatts) suggest that significant mass savings may be realized by implementing a direct-drive power system, so NASA recently established the National Direct-Drive Testbed to examine technical issues identified by previous investigations. The testbed includes a 12-kW solar array and power control station designed to power single and multiple Hall thrusters over a wide range of voltages and currents. In this paper, single Hall thruster operation directly from solar array output at discharge voltages of 200 to 450 V and discharge powers of 1 to 10 kW is reported. Hall thruster control and operation is shown to be simple and no different than for operation on conventional power supplies. Thruster and power system electrical oscillations were investigated over a large range of operating conditions and with different filter capacitances. Thruster oscillations were the same as for conventional power supplies, did not adversely affect solar array operation, and were independent of filter capacitance from 8 to 80 ?F. Solar array current and voltage oscillations were very small compared to their mean values and showed a modest dependence on capacitor size. No instabilities or anomalous behavior were observed in the thruster or power system at any operating condition investigated, including near and at the array peak power point. Thruster startup using the anode propellant flow as the power 'switch' was shown to be simple and reliable with system transients mitigated by the proper selection of filter capacitance size. Shutdown via cutoff of propellant flow was also demonstrated. A simple electrical circuit model was developed and is shown to have good agreement with the experimental data.

In order to reveal the internal phenomenon theoretically within the electron cyclotron resonance ion thruster μ 10, internal microwave electric field measurement is very important because it is closely related to plasma producing mechanism. We have established a technology of electric field measurement with an optical fiber sensor which uses an Electro-Optic crystal (EO probe). This technology enables electric field measurement in plasma source under beam acceleration without disturbing microwave electric field. In this study, first, validity of electric field measurement using the EO probe in the atmosphere was demonstrated by comparing experimental results with FDTD simulation. Then, we measured axial electric field distribution in the accelerated plasma. This experiments indicated that electric field distribution in the μ10 thruster was related to its beam current .

Summary form only given. The effect of inductive coil geometry on the operating characteristics of an inductive pulsed plasma thruster [1,2] is investigated analytically and experimentally. Model results indicate that the introduction of radial current sheet motion caused by a conical inductive coil geometry (versus a flat circular plate) increases the axial dynamic impedance parameter at which thrust efficiency is maximized and generally decreases the overall achievable thrust efficiency. Operational characteristics of two thrusters with inductive coils of different cone angles are explored through thrust stand measurements and time-integrated, unfiltered photography. Trends in impulse bit measurements indicate that, in the present configuration, the thruster with the inductive coil possessing a smaller cone angle produced larger values of thrust, in apparent contradiction to results of the model.

Summary form only given. A radiofrequency plasma source was previously developed at the University of Michigan to study the behavior of ions downstream of a diverging magnetic field.1,2 A follow-up experiment investigates the performance of the plasma source operating as a thruster. During this experiment, simultaneous thruster performance and far-field plume measurements of the rf plasma source are taken in the Large Vacuum Test Facility (LVTF) at the University of Michigan. Thruster performance is measured using an inverted-pendulum thrust stand, and plume properties are investigated using a nude Faraday probe, a retarding potential analyzer, and a commercial rf-compensated Langmuir probe. The pressure in the LVTF is held below 5.5×10−6 torr during operation of the thruster in order to reduce facility effects on probe measurements. The work presented here focuses on the results of the probe analysis.

Summary form only given. Recently a new device namely micro-cathode arc thruster (μCAT) based on vacuum arc discharge has been developed for the spacecraft propulsion application1. The μCAT is a micro-Newton level thrust device that employs a magnetic field and is suited for the long-time operation of micro-thruster for the nano-satellite propulsion such as Cubesats2, 3.

Squid and jellyfish generate propulsive forces by successively taking in and expelling high momentum jets of water. This method of propulsion offers several advantages to underwater vehicles/robots. The driving mechanism can be placed internal to the vehicle, reducing the drag associated with an abundance of external thrusters and control surfaces. The thrusters can generate accurate predictable forcing in the low thrust range, while still generating the thrust nearly instantaneously over the entire force range. Vortex ring formation dynamics play a key role in creating thrust. It is observed that squid and jellyfish eject fluid jets which are not exactly parallel, and have a contracting velocity in the radial direction. We generated jet flows which were both parallel, and converging and directly measured the thrust acting on the device, as well as the total impulse of the jet, to determine the effect a non-zero radial velocity had on thrust production. The radial velocity was observed to increase the total impulse of the jet by 60% for large stroke ratio jets.

International audience; A global (volume-averaged) model of a gridded-ion thruster is proposed. The neutral propellant (xenon gas) is injected into the thruster chamber at a fixed rate and a plasma is generated by circulating a radiofrequency current in an inductive coil. The ions generated in this plasma are accelerated out of the thruster by a pair of DC biased grids. The neutralization downstream is not treated. Xenon atoms also flow out of the thruster across the grids. The model, based on particle and energy balance equations, solves for four global variables in the thruster chamber: the plasma density, the electron temperature, the neutral gas (atom) density, and the neutral gas temperature. The important quantities to evaluate the thruster efficiency and performances are calculated from these variables and from the voltage across the grids. It is found that the mass utilization efficiency rapidly decreases with the gas flow rate. However, the radiofrequency power transfer efficiency increases significantly with the injected gas flow rate. Therefore, there is a compromise to be found between these two quantities.

Local, time‐dependent magnetic field measurements have been made in the Los Alamos coaxial thruster experiment (CTX) [C. W. Barnes et al., Phys. Fluids B 2, 1871 (1990); J. C. Fernandez et al., Nucl. Fusion 28, 1555 (1988)] using a 24 coil magnetic probe array (eight spatial positions, three axis probes). The CTX is a magnetized, coaxial plasma gun presently being used to investigate the viability of high pulsed power plasma thrusters for advanced electric propulsion. Previous efforts on this device have indicated that high pulsed power plasma guns are attractive candidates for advanced propulsion that employ ideal magnetohydrodynamic (MHD) plasma stream flow through self‐formed magnetic nozzles. Indirect evidence of magnetic nozzle formation was obtained from plasma gun performance and measurements of directed axial velocities up to vz∼107 cm/s. The purpose of this work is to make direct measurement of the time evolving magnetic field topology. The intent is to both identify that applied magnetic field d...

The Los Alamos National Laboratory Coaxial Thruster Experiment (CTX) has been upgraded to enable 10 ms quasi-steady-state (QSS) operation of magnetohydrodynamic type thrusters at power levels from 1 to 40 MW. Here we report on experimental observations and associated analysis for QSS discharges using deuterium, helium, and argon propellants. Measurements of thruster current and voltage, magnetic field fluctuations, electron density and temperature, and plasma potential indicate true quasi-steady operation over 10 ms with propellant exhaust velocities of 1-2/spl times/10/sup 5/ m/s in helium and deuterium. The application of unique applied magnetic nozzle configurations result in a substantial reduction of the measured anode fall potential. Data analysis and analytical MHD modeling imply an impulsive start in the propellant flow that is limited to the Alfven critical ionization velocity and the presence of an effective constriction or "nozzling" in the how channel. This research indicates that efficiency optimization of MHD class thrusters will most likely occur at megawatt power levels. This implies repetitively pulsed, high-power operation for near term, low-mean-power solar electric propulsion missions. >

With the growing interest in Hall thruster technology, comes the need to fully characterize the plasma dynamics that determine performance. Of particular interest is the existence of a periodic low frequency oscillation, commonly referred to as the breathing mode. Although there has been a significant effort to quantify plasma properties, this has been primarily limited to time-averaged measurements. In order to fully understand the periodic instabilities characteristic of Hall thruster behavior, time resolved techniques must be developed. This study presents a non-intrusive method of determining time-resolved electron temperature fluctuations during the breathing mode cycle. A triggering system was developed to synchronize measurements to thruster discharge current oscillations. Emission in the NIR is measured near the thruster exit plane using an optical fiber coupled to a 1.25 m focal length spectrometer and intensified CCD (ICCD) detector. ICCD gating is controlled using the output of the triggering system and a user controlled gate delay. This allows emission sampling only from the desired portion of the breathing mode cycle. Emission is integrated over several periods to allow for improved signal to noise ratio (SNR). Electron temperature is determined using the intensity of Xe I emission lines in the context of a collisional-radiative model. The resulting measured electron temperature fluctuations during a nominal breathing mode cycle are presented for a 600 W Hall effect thruster at a single nominal operating condition.