Mode transition of a Hall thruster discharge plasma

International Nuclear Information System (INIS)

Hara, Kentaro; Sekerak, Michael J.; Boyd, Iain D.; Gallimore, Alec D.

2014-01-01

A Hall thruster is a cross-field plasma device used for spacecraft propulsion. An important unresolved issue in the development of Hall thrusters concerns the effect of discharge oscillations in the range of 10–30 kHz on their performance. The use of a high speed Langmuir probe system and ultra-fast imaging of the discharge plasma of a Hall thruster suggests that the discharge oscillation mode, often called the breathing mode, is strongly correlated to an axial global ionization mode. Stabilization of the global oscillation mode is achieved as the magnetic field is increased and azimuthally rotating spokes are observed. A hybrid-direct kinetic simulation that takes into account the transport of electronically excited atoms is used to model the discharge plasma of a Hall thruster. The predicted mode transition agrees with experiments in terms of the mean discharge current, the amplitude of discharge current oscillation, and the breathing mode frequency. It is observed that the stabilization of the global oscillation mode is associated with reduced electron transport that suppresses the ionization process inside the channel. As the Joule heating balances the other loss terms including the effects of wall loss and inelastic collisions, the ionization oscillation is damped, and the discharge oscillation stabilizes. A wide range of the stable operation is supported by the formation of a space charge saturated sheath that stabilizes the electron axial drift and balances the Joule heating as the magnetic field increases. Finally, it is indicated from the numerical results that there is a strong correlation between the emitted light intensity and the discharge current.

Electrostatic ion thrusters - towards predictive modeling

Energy Technology Data Exchange (ETDEWEB)

Kalentev, O.; Matyash, K.; Duras, J.; Lueskow, K.F.; Schneider, R. [Ernst-Moritz-Arndt Universitaet Greifswald, D-17489 (Germany); Koch, N. [Technische Hochschule Nuernberg Georg Simon Ohm, Kesslerplatz 12, D-90489 Nuernberg (Germany); Schirra, M. [Thales Electronic Systems GmbH, Soeflinger Strasse 100, D-89077 Ulm (Germany)

2014-02-15

The development of electrostatic ion thrusters so far has mainly been based on empirical and qualitative know-how, and on evolutionary iteration steps. This resulted in considerable effort regarding prototype design, construction and testing and therefore in significant development and qualification costs and high time demands. For future developments it is anticipated to implement simulation tools which allow for quantitative prediction of ion thruster performance, long-term behavior and space craft interaction prior to hardware design and construction. Based on integrated numerical models combining self-consistent kinetic plasma models with plasma-wall interaction modules a new quality in the description of electrostatic thrusters can be reached. These open the perspective for predictive modeling in this field. This paper reviews the application of a set of predictive numerical modeling tools on an ion thruster model of the HEMP-T (High Efficiency Multi-stage Plasma Thruster) type patented by Thales Electron Devices GmbH. (copyright 2014 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim) (orig.)

Particle simulation of grid system for krypton ion thrusters

Directory of Open Access Journals (Sweden)

Maolin CHEN

2018-04-01

Full Text Available The transport processes of plasmas in grid systems of krypton (Kr ion thrusters at different acceleration voltages were simulated with a 3D-PIC model, and the result was compared with xenon (Xe ion thrusters. The variation of the screen grid transparency, the accelerator grid current ratio and the divergence loss were explored. It is found that the screen grid transparency increases with the acceleration voltage and decreases with the beam current, while the accelerator grid current ratio and divergence loss decrease first and then increase with the beam current. This result is the same with Xe ion thrusters. Simulation results also show that Kr ion thrusters have more advantages than Xe ion thrusters, such as higher screen grid transparency, smaller accelerator grid current ratio, larger cut-off current threshold, and better divergence loss characteristic. These advantages mean that Kr ion thrusters have the ability of operating in a wide range of current. Through comprehensive analyses, it can be concluded that using Kr as propellant is very suitable for a multi-mode ion thruster design. Keywords: Grid system, Ion thrusters, Krypton, Particle in cell method, Plasma

Cathode Effects in Cylindrical Hall Thrusters

Energy Technology Data Exchange (ETDEWEB)

Granstedt, E.M.; Raitses, Y.; Fisch, N. J.

2008-09-12

Stable operation of a cylindrical Hall thruster (CHT) has been achieved using a hot wire cathode, which functions as a controllable electron emission source. It is shown that as the electron emission from the cathode increases with wire heating, the discharge current increases, the plasma plume angle reduces, and the ion energy distribution function shifts toward higher energies. The observed effect of cathode electron emission on thruster parameters extends and clarifies performance improvements previously obtained for the overrun discharge current regime of the same type of thruster, but using a hollow cathode-neutralizer. Once thruster discharge current saturates with wire heating, further filament heating does not affect other discharge parameters. The saturated values of thruster discharge parameters can be further enhanced by optimal placement of the cathode wire with respect to the magnetic field.

Magnetoelectrostatic thruster physical geometry tests

Science.gov (United States)

Ramsey, W. D.

1981-01-01

Inert gas tests are conducted with several magnetoelectrostatic containment discharge chamber geometries. The configurations tested include three discharge chamber lengths; three boundary magnet patterns; two different flux density magnet materials; hemispherical and conical shaped thrusters having different surface-to-volume ratios; and two and three grid ion optics. Argon mass utilizations of 60 to 79% are attained at 210 to 280 eV/ion in different test configurations. Short hemi thruster configurations are found to produce 70 to 92% xenon mass utilization at 185 to 220 eV/ion.

15 cm mercury multipole thruster

Science.gov (United States)

Longhurst, G. R.; Wilbur, P. J.

1978-01-01

A 15 cm multipole ion thruster was adapted for use with mercury propellant. During the optimization process three separable functions of magnetic fields within the discharge chamber were identified: (1) they define the region where the bulk of ionization takes place, (2) they influence the magnitudes and gradients in plasma properties in this region, and (3) they control impedance between the cathode and main discharge plasmas in hollow cathode thrusters. The mechanisms for these functions are discussed. Data from SERT II and cusped magnetic field thrusters are compared with those measured in the multipole thruster. The performance of this thruster is shown to be similar to that of the other two thrusters. Means of achieving further improvement in the performance of the multipole thruster are suggested.

Parametric Investigation of Miniaturized Cylindrical and Annular Hall Thrusters

International Nuclear Information System (INIS)

Smirnov, A.; Raitses, Y.; Fisch, N.J.

2002-01-01

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

Krypton Ion Thruster Performance

Science.gov (United States)

Patterson, Michael J.; Williams, George J.

1992-01-01

Preliminary data were obtained from a 30 cm ion thruster operating on krypton propellant over the input power range of 0.4 to 5.5 kW. The data presented are compared and contrasted to the data obtained with xenon propellant over the same input power envelope. Typical krypton thruster efficiency was 70 percent at a specific impulse of approximately 5000 s, with a maximum demonstrated thrust to power ratio of approximately 42 mN/kW at 2090 s specific impulse and 1580 watts input power. Critical thruster performance and component lifetime issues were evaluated. Order of magnitude power throttling was demonstrated using a simplified power-throttling strategy.

NSTAR Ion Thruster and Breadboard Power Processor Functional Integration Test Results

Science.gov (United States)

Hamley, John A.; Pinero, Luis R.; Rawlin, Vincent K.; Miller, John R.; Myers, Roger M.; Bowers, Glen E.

1996-01-01

A 2.3 kW Breadboard Power Processing Unit (BBPPU) was developed as part of the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) Program. The NSTAR program will deliver an electric propulsion system based on a 30 cm xenon ion thruster to the New Millennium (NM) program for use as the primary propulsion system for the initial NM flight. The final development test for the BBPPU, the Functional Integration Test, was carried out to demonstrate all aspects of BBPPU operation with an Engineering Model Thruster. Test objectives included: (1) demonstration and validation of automated thruster start procedures, (2) demonstration of stable closed loop control of the thruster beam current, (3) successful response and recovery to thruster faults, and (4) successful safing of the system during simulated spacecraft faults. These objectives were met over the specified 80-120 VDC input voltage range and 0.5-2.3 output power capability of the BBPPU. Two minor anomalies were noted in discharge and neutralizer keeper current. These anomalies did not affect the stability of the system and were successfully corrected.

HG ion thruster component testing

Science.gov (United States)

Mantenieks, M. A.

1979-01-01

Cathodes, isolators, and vaporizers are critical components in determining the performance and lifetime of mercury ion thrusters. The results of life tests of several of these components are reported. A 30-cm thruster CIV test in a bell jar has successfully accumulated over 26,000 hours. The cathode has undergone 65 restarts during the life test without requiring any appreciable increases in starting power. Recently, all restarts have been achieved with only the 44 volt keeper supply with no change required in the starting power. Another ongoing 30-cm Hg thruster cathode test has successfully passed the 10,000 hour mark. A solid-insert, 8-cm thruster cathode has accumulated over 4,000 hours of thruster operation. All starts have been achieved without the use of a high voltage ignitor. The results of this test indicate that the solid impregnated insert is a viable neutralizer cathode for the 8-cm thruster.

Optimization of Cylindrical Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Yevgeny; Smirnov, Artem; Granstedt, Erik; Fisch, Nathaniel J.

2007-01-01

The cylindrical Hall thruster features high ionization efficiency, quiet operation, and ion acceleration in a large volume-to-surface ratio channel with performance comparable with the state-of-the-art annular Hall thrusters. These characteristics were demonstrated in low and medium power ranges. Optimization of miniaturized cylindrical thrusters led to performance improvements in the 50-200W input power range, including plume narrowing, increased thruster efficiency, reliable discharge initiation, and stable operation.

Optimization of Cylindrical Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Yevgeny; Smirnov, Artem; Granstedt, Erik; Fi, Nathaniel J.

2007-01-01

The cylindrical Hall thruster features high ionization efficiency, quiet operation, and ion acceleration in a large volume-to-surface ratio channel with performance comparable with the state-of-the-art annular Hall thrusters. These characteristics were demonstrated in low and medium power ranges. Optimization of miniaturized cylindrical thrusters led to performance improvements in the 50-200W input power range, including plume narrowing, increased thruster efficiency, reliable discharge initiation, and stable operation

Discharge Oscillations in a Permanent Magnet Cylindrical Hall-Effect Thruster

Science.gov (United States)

Polzin, K. A.; Sooby, E. S.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

Measurements of the discharge current in a cylindrical Hall thruster are presented to quantify plasma oscillations and instabilities without introducing an intrusive probe into the plasma. The time-varying component of the discharge current is measured using a current monitor that possesses a wide frequency bandwidth and the signal is Fourier transformed to yield the frequency spectra present, allowing for the identification of plasma oscillations. The data show that the discharge current oscillations become generally greater in amplitude and complexity as the voltage is increased, and are reduced in severity with increasing flow rate. The breathing mode ionization instability is identified, with frequency as a function of discharge voltage not increasing with discharge voltage as has been observed in some traditional Hall thruster geometries, but instead following a scaling similar to a large-amplitude, nonlinear oscillation mode recently predicted in for annular Hall thrusters. A transition from lower amplitude oscillations to large relative fluctuations in the oscillating discharge current is observed at low flow rates and is suppressed as the mass flow rate is increased. A second set of peaks in the frequency spectra are observed at the highest propellant flow rate tested. Possible mechanisms that might give rise to these peaks include ionization instabilities and interactions between various oscillatory modes.

Experimental Studies of Anode Sheath Phenomena in a Hall Thruster Discharge

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.

2004-01-01

Both electron-repelling and electron-attracting anode sheaths in a Hall thruster were characterized by measuring the plasma potential with biased and emissive probes [L. Dorf, Y. Raitses, V. Semenov, and N.J. Fisch, Appl. Phys. Let. 84 (2004) 1070]. In the present work, two-dimensional structures of the plasma potential, electron temperature, and plasma density in the near-anode region of a Hall thruster with clean and dielectrically coated anodes are identified. Possible mechanisms of anode sheath formation in a Hall thruster are analyzed. The path for current closure to the anode appears to be the determining factor in the anode sheath formation process. The main conclusion of this work is that the anode sheath formation in Hall thrusters differs essentially from that in the other gas discharge devices, like a glow discharge or a hollow anode, because the Hall thruster utilizes long electron residence times to ionize rather than high neutral pressures

High-Power Ion Thruster Technology

Science.gov (United States)

Beattie, J. R.; Matossian, J. N.

1996-01-01

Performance data are presented for the NASA/Hughes 30-cm-diam 'common' thruster operated over the power range from 600 W to 4.6 kW. At the 4.6-kW power level, the thruster produces 172 mN of thrust at a specific impulse of just under 4000 s. Xenon pressure and temperature measurements are presented for a 6.4-mm-diam hollow cathode operated at emission currents ranging from 5 to 30 A and flow rates of 4 sccm and 8 sccm. Highly reproducible results show that the cathode temperature is a linear function of emission current, ranging from approx. 1000 C to 1150 C over this same current range. Laser-induced fluorescence (LIF) measurements obtained from a 30-cm-diam thruster are presented, suggesting that LIF could be a valuable diagnostic for real-time assessment of accelerator-arid erosion. Calibration results of laminar-thin-film (LTF) erosion badges with bulk molybdenum are presented for 300-eV xenon, krypton, and argon sputtering ions. Facility-pressure effects on the charge-exchange ion current collected by 8-cm-diam and 30-cm-diam thrusters operated on xenon propellant are presented to show that accel current is nearly independent of facility pressure at low pressures, but increases rapidly under high-background-pressure conditions.

Power processing systems for ion thrusters.

Science.gov (United States)

Herron, B. G.; Garth, D. R.; Finke, R. C.; Shumaker, H. A.

1972-01-01

The proposed use of ion thrusters to fulfill various communication satellite propulsion functions such as east-west and north-south stationkeeping, attitude control, station relocation and orbit raising, naturally leads to the requirement for lightweight, efficient and reliable thruster power processing systems. Collectively, the propulsion requirements dictate a wide range of thruster power levels and operational lifetimes, which must be matched by the power processing. This paper will discuss the status of such power processing systems, present system design alternatives and project expected near future power system performance.

Status of the J-series 30-cm mercury ion thruster

Science.gov (United States)

Kami, S.; Dulgeroff, C. R.; Bechtel, R. T.

1982-01-01

This paper describes the status of the 30-cm J-series mercury ion thruster. This thruster was baselined for the Solar Electric Propulsion System (SEPS) vehicle. This thruster is described and several modifications plus suggested modifications are presented. Some of the modifications resulted from tests performed with the thruster. The operational characteristics of eight J-series thrusters are presented. Isolator contamination and flake formation are also discussed.

Brayton-Cycle Power-Conversion Unit Tested With Ion Thruster

Science.gov (United States)

Hervol, David S.

2005-01-01

Nuclear electric propulsion has been identified as an enabling technology for future NASA space science missions, such as the Jupiter Icy Moons Orbiter (JIMO) now under study. An important element of the nuclear electric propulsion spacecraft is the power conversion system, which converts the reactor heat to electrical power for use by the ion propulsion system and other spacecraft loads. The electrical integration of the power converter and ion thruster represents a key technical challenge in making nuclear electric propulsion technology possible. This technical hurdle was addressed extensively on December 1, 2003, when a closed- Brayton-cycle power-conversion unit was tested with a gridded ion thruster at the NASA Glenn Research Center. The test demonstrated end-to-end power throughput and marked the first-ever coupling of a Brayton turbo alternator and a gridded ion thruster, both of which are candidates for use on JIMO-type missions. The testing was conducted at Glenn's Vacuum Facility 6, where the Brayton unit was installed in the 3-m-diameter vacuum test port and the ion thruster was installed in the 7.6-m-diameter main chamber.

Cylindrical Hall Thrusters with Permanent Magnets

International Nuclear Information System (INIS)

Raitses, Yevgeny; Merino, Enrique; Fisch, Nathaniel J.

2010-01-01

The use of permanent magnets instead of electromagnet coils for low power Hall thrusters can offer a significant reduction of both the total electric power consumption and the thruster mass. Two permanent magnet versions of the miniaturized cylindrical Hall thruster (CHT) of different overall dimensions were operated in the power range of 50W-300 W. The discharge and plasma plume measurements revealed that the CHT thrusters with permanent magnets and electromagnet coils operate rather differently. In particular, the angular ion current density distribution from the permanent magnet thrusters has an unusual halo shape, with a majority of high energy ions flowing at large angles with respect to the thruster centerline. Differences in the magnetic field topology outside the thruster channel and in the vicinity of the channel exit are likely responsible for the differences in the plume characteristics measured for the CHTs with electromagnets and permanent magnets. It is shown that the presence of the reversing-direction or cusp-type magnetic field configuration inside the thruster channel without a strong axial magnetic field outside the thruster channel does not lead to the halo plasma plume from the CHT.

Post-Test Inspection of NASA's Evolutionary Xenon Thruster Long-Duration Test Hardware: Discharge and Neutralizer Cathodes

Science.gov (United States)

Shastry, Rohit; Soulas, George C.

2016-01-01

The NEXT Long-Duration Test is part of a comprehensive thruster service life assessment intended to demonstrate overall throughput capability, validate service life models, quantify wear rates as a function of time and operating condition, and identify any unknown life-limiting mechanisms. The test was voluntarily terminated in February 2014 after demonstrating 51,184 hours of high-voltage operation, 918 kg of propellant throughput, and 35.5 MN-s of total impulse. The post-test inspection of the thruster hardware began shortly afterwards with a combination of non-destructive and destructive analysis techniques, and is presently nearing completion. This paper presents relevant results of the post-test inspection for both discharge and neutralizer cathodes. Discharge keeper erosion was found to be significantly reduced from what was observed in the NEXT 2 kh wear test and NSTAR Extended Life Test, providing adequate protection of vital cathode components throughout the test with ample lifetime remaining. The area of the discharge cathode orifice plate that was exposed by the keeper orifice exhibited net erosion, leading to cathode plate material building up in the cathode-keeper gap and causing a thermally-induced electrical short observed during the test. Significant erosion of the neutralizer cathode orifice was also found and is believed to be the root cause of an observed loss in flow margin. Deposition within the neutralizer keeper orifice as well as on the downstream surface was thicker than expected, potentially resulting in a facility-induced impact on the measured flow margin from plume mode. Neutralizer keeper wall erosion on the beam side was found to be significantly lower compared to the NEXT 2 kh wear test, likely due to the reduction in beam extraction diameter of the ion optics that resulted in decreased ion impingement. Results from the post-test inspection have led to some minor thruster design improvements.

Electron energy distribution function in a low-power Hall thruster discharge and near-field plume

Science.gov (United States)

Tichý, M.; Pétin, A.; Kudrna, P.; Horký, M.; Mazouffre, S.

2018-06-01

Electron temperature and plasma density, as well as the electron energy distribution function (EEDF), have been obtained inside and outside the dielectric channel of a 200 W permanent magnet Hall thruster. Measurements were carried out by means of a cylindrical Langmuir probe mounted onto a compact fast moving translation stage. The 3D particle-in cell numerical simulations complement experiments. The model accounts for the crossed electric and magnetic field configuration in a weakly collisional regime where only electrons are magnetized. Since only the electron dynamics is of interest in this study, an artificial mass of ions corresponding to mi = 30 000me was used to ensure ions could be assumed at rest. The simulation domain is located at the thruster exit plane and does not include the cathode. The measured EEDF evidences a high-energy electron population that is superimposed onto the low energy bulk population outside the channel. Inside the channel, the EEDF is close to Maxwellian. Both the experimental and numerical EEDF depart from an equilibrium distribution at the channel exit plane, a region of high magnetic field. We therefore conclude that the fast electron group found in the experiment corresponds to the electrons emitted by the external cathode that reach the thruster discharge without experiencing collision events.

Theoretical and experimental study of a thruster discharging a weight

Science.gov (United States)

Michaels, Dan; Gany, Alon

2014-06-01

An innovative concept for a rocket type thruster that can be beneficial for spacecraft trajectory corrections and station keeping was investigated both experimentally and theoretically. It may also be useful for divert and attitude control systems (DACS). The thruster is based on a combustion chamber discharging a weight through an exhaust tube. Calculations with granular double-base propellant and a solid ejected weight reveal that a specific impulse based on the propellant mass of well above 400 s can be obtained. An experimental thruster was built in order to demonstrate the new idea and validate the model. The thruster impulse was measured both directly with a load cell and indirectly by using a pressure transducer and high speed photography of the weight as it exits the tube, with both ways producing very similar total impulse measurement. The good correspondence between the computations and the measured data validates the model as a useful tool for studying and designing such a thruster.

High Input Voltage Hall Thruster Discharge Converter, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — The overall scope of this Phase I/II effort is the development of a high efficiency 15kW (nominal) Hall thruster discharge converter. In Phase I, Busek Co. Inc. will...

Transit-time instability in Hall thrusters

International Nuclear Information System (INIS)

Barral, Serge; Makowski, Karol; Peradzynski, Zbigniew; Dudeck, Michel

2005-01-01

Longitudinal waves characterized by a phase velocity of the order of the velocity of ions have been recurrently observed in Hall thruster experiments and simulations. The origin of this so-called ion transit-time instability is investigated with a simple one-dimensional fluid model of a Hall thruster discharge in which cold ions are accelerated between two electrodes within a quasineutral plasma. A short-wave asymptotics applied to linearized equations shows that plasma perturbations in such a device consist of quasineutral ion acoustic waves superimposed on a background standing wave generated by discharge current oscillations. Under adequate circumstances and, in particular, at high ionization levels, acoustic waves are amplified as they propagate, inducing strong perturbation of the ion density and velocity. Responding to the subsequent perturbation of the column resistivity, the discharge current generates a standing wave, the reflection of which sustains the generation of acoustic waves at the inlet boundary. A calculation of the frequency and growth rate of this resonance mechanism for a supersonic ion flow is proposed, which illustrates the influence of the ionization degree on their onset and the approximate scaling of the frequency with the ion transit time. Consistent with experimental reports, the traveling wave can be observed on plasma density and velocity perturbations, while the plasma potential ostensibly oscillates in phase along the discharge

Effect of Anode Magnetic Shield on Magnetic Field and Ion Beam in Cylindrical Hall Thruster

International Nuclear Information System (INIS)

Zhao Jie; Wang Shiqing; Liu Jian; Xu Li; Tang Deli; Geng Shaofei

2010-01-01

Numerical simulation of the effect of the anode magnetic shielding on the magnetic field and ion beam in a cylindrical Hall thruster is presented. The results show that after the anode is shielded by the magnetic shield, the magnetic field lines near the anode surface are obviously convex curved, the ratio of the magnetic mirror is enhanced, the width of the positive magnetic field gradient becomes larger than that without the anode magnetic shielding, the radial magnetic field component is enhanced, and the discharge plasma turbulence is reduced as a result of keeping the original saddle field profile and the important role the other two saddle field profiles play in restricting electrons. The results of the particle in cell (PIC) numerical simulation show that both the ion number and the energy of the ion beam increase after the anode is shielded by the magnetic shield. In other words, the specific impulse of the cylindrical Hall thruster is enhanced.

Thermal Environmental Testing of NSTAR Engineering Model Ion Thrusters

Science.gov (United States)

Rawlin, Vincent K.; Patterson, Michael J.; Becker, Raymond A.

1999-01-01

NASA's New Millenium program will fly a xenon ion propulsion system on the Deep Space 1 Mission. Tests were conducted under NASA's Solar Electric Propulsion Technology Applications Readiness (NSTAR) Program with 3 different engineering model ion thrusters to determine thruster thermal characteristics over the NSTAR operating range in a variety of thermal environments. A liquid nitrogen-cooled shroud was used to cold-soak the thruster to -120 C. Initial tests were performed prior to a mature spacecraft design. Those results and the final, severe, requirements mandated by the spacecraft led to several changes to the basic thermal design. These changes were incorporated into a final design and tested over a wide range of environmental conditions.

Thermo-mechanical design aspects of mercury bombardment ion thrusters.

Science.gov (United States)

Schnelker, D. E.; Kami, S.

1972-01-01

The mechanical design criteria are presented as background considerations for solving problems associated with the thermomechanical design of mercury ion bombardment thrusters. Various analytical procedures are used to aid in the development of thruster subassemblies and components in the fields of heat transfer, vibration, and stress analysis. Examples of these techniques which provide computer solutions to predict and control stress levels encountered during launch and operation of thruster systems are discussed. Computer models of specific examples are presented.

Retrofit and verification test of a 30-cm ion thruster

Science.gov (United States)

Dulgeroff, C. R.; Poeschel, R. L.

1980-01-01

Twenty modifications were found to be necessary and were approved by design review. These design modifications were incorporated in the thruster documents (drawings and procedures) to define the J series thruster. Sixteen of the design revisions were implemented in a 900 series thruster by retrofit modification. A standardized set of test procedures was formulated, and the retrofit J series thruster design was verified by test. Some difficulty was observed with the modification to the ion optics assembly, but the overall effect of the design modification satisfies the design objectives. The thruster was tested over a wide range of operating parameters to demonstrate its capabilities.

Plasma Perturbations in High-Speed Probing of Hall Thruster Discharge Chambers: Quantification and Mitigation

Science.gov (United States)

Jorns, Benjamin A.; Goebel, Dan M.; Hofer, Richard R.

2015-01-01

An experimental investigation is presented to quantify the effect of high-speed probing on the plasma parameters inside the discharge chamber of a 6-kW Hall thruster. Understanding the nature of these perturbations is of significant interest given the importance of accurate plasma measurements for characterizing thruster operation. An array of diagnostics including a high-speed camera and embedded wall probes is employed to examine in real time the changes in electron temperature and plasma potential induced by inserting a high-speed reciprocating Langmuir probe into the discharge chamber. It is found that the perturbations onset when the scanning probe is downstream of the electron temperature peak, and that along channel centerline, the perturbations are best characterized as a downstream shift of plasma parameters by 15-20% the length of the discharge chamber. A parametric study is performed to investigate techniques to mitigate the observed probe perturbations including varying probe speed, probe location, and operating conditions. It is found that the perturbations largely disappear when the thruster is operated at low power and low discharge voltage. The results of this mitigation study are discussed in the context of recommended methods for generating unperturbed measurements of the discharge chamber plasma.

Recycle Requirements for NASA's 30 cm Xenon Ion Thruster

Science.gov (United States)

Pinero, Luis R.; Rawlin, Vincent K.

1994-01-01

Electrical breakdowns have been observed during ion thruster operation. These breakdowns, or arcs, can be caused by several conditions. In flight systems, the power processing unit must be designed to handle these faults autonomously. This has a strong impact on power processor requirements and must be understood fully for the power processing unit being designed for the NASA Solar Electric Propulsion Technology Application Readiness program. In this study, fault conditions were investigated using a NASA 30 cm ion thruster and a power console. Power processing unit output specifications were defined based on the breakdown phenomena identified and characterized.

Effects of magnetic field strength in the discharge channel on the performance of a multi-cusped field thruster

Directory of Open Access Journals (Sweden)

Peng Hu

2016-09-01

Full Text Available 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.

Pocket rocket: An electrothermal plasma micro-thruster

Science.gov (United States)

Greig, Amelia Diane

Recently, an increase in use of micro-satellites constructed from commercial off the shelf (COTS) components has developed, to address the large costs associated with designing, testing and launching satellites. One particular type of micro-satellite of interest are CubeSats, which are modular 10 cm cubic satellites with total weight less than 1.33 kg. To assist with orbit boosting and attitude control of CubeSats, micro-propulsion systems are required, but are currently limited. A potential electrothermal plasma micro-thruster for use with CubeSats or other micro-satellites is under development at The Australian National University and forms the basis for this work. The thruster, known as ‘Pocket Rocket’, utilises neutral gas heating from ion-neutral collisions within a weakly ionised asymmetric plasma discharge, increasing the exhaust thermal velocity of the propellant gas, thereby producing higher thrust than if the propellant was emitted cold. In this work, neutral gas temperature of the Pocket Rocket discharge is studied in depth using rovibrational spectroscopy of the nitrogen (N2) second positive system (C3Πu → B3Πg), using both pure N2 and argon/N2 mixtures as the operating gas. Volume averaged steady state gas temperatures are measured for a range of operating conditions, with an analytical collisional model developed to verify experimental results. Results show that neutral gas heating is occurring with volume averaged steady state temperatures reaching 430 K in N2 and 1060 K for argon with 1% N2 at standard operating conditions of 1.5 Torr pressure and 10 W power input, demonstrating proof of concept for the Pocket Rocket thruster. Spatiotemporal profiles of gas temperature identify that the dominant heating mechanisms are ion-neutral collisions within the discharge and wall heating from ion bombardment of the thruster walls. To complement the experimental results, computational fluid dynamics (CFD) simulations using the commercial CFD

Performance of an iodine-fueled radio-frequency ion-thruster

Science.gov (United States)

Holste, Kristof; Gärtner, Waldemar; Zschätzsch, Daniel; Scharmann, Steffen; Köhler, Peter; Dietz, Patrick; Klar, Peter J.

2018-01-01

Two sets of performance data of the same radio-frequency ion-thruster (RIT) have been recorded using iodine and xenon, respectively, as propellant. To characterize the thruster's performance, we have recorded the radio-frequency DC-power, required for yielding preset values of the extracted ion-beam currents, as a function of mass flow. For that purpose, an iodine mass flow system had to be developed, calibrated, and integrated into a newly-built test facility for studying corrosive propellants. The performance mappings for iodine and xenon differ significantly despite comparable operation conditions. At low mass flows, iodine exhibits the better performance. The situation changes at higher mass flows where the performance of iodine is significantly poorer than that of xenon. The reason is very likely related to the molecular nature of iodine. Our results show that iodine as propellant is compatible with RIT technology. Furthermore, it is a viable alternative as propellant for dedicated space missions. In particular, when taking into account additional benefits such as possible storage as a solid and its low price the use of iodine as propellant in ion thrusters is competitive.

Numerical research of a 2D axial symmetry hybrid model for the radio-frequency ion thruster

Science.gov (United States)

Chenchen, WU; Xinfeng, SUN; Zuo, GU; Yanhui, JIA

2018-04-01

Since the high efficiency discharge is critical to the radio-frequency ion thruster (RIT), a 2D axial symmetry hybrid model has been developed to study the plasma evolution of RIT. The fluid method and the drift energy correction of the electron energy distribution function (EEDF) are applied to the analysis of the RIT discharge. In the meantime, the PIC-MCC method is used to investigate the ion beam current extraction character for the plasma plume region. The beam current simulation results, with the hybrid model, agree well with the experimental results, and the error is lower than 11%, which shows the validity of the model. The further study shows there is an optimal ratio for the radio-frequency (RF) power and the beam current extraction power under the fixed RIT configuration. And the beam extraction efficiency will decrease when the discharge efficiency beyond a certain threshold (about 87 W). As the input parameters of the hybrid model are all the design values, it can be directly used to the optimum design for other kinds of RITs and radio-frequency ion sources.

Integration Testing of a Modular Discharge Supply for NASA's High Voltage Hall Accelerator Thruster

Science.gov (United States)

Pinero, Luis R.; Kamhawi, hani; Drummond, Geoff

2010-01-01

NASA s In-Space Propulsion Technology Program is developing a high performance Hall thruster that can fulfill the needs of future Discovery-class missions. The result of this effort is the High Voltage Hall Accelerator thruster that can operate over a power range from 0.3 to 3.5 kW and a specific impulse from 1,000 to 2,800 sec, and process 300 kg of xenon propellant. Simultaneously, a 4.0 kW discharge power supply comprised of two parallel modules was developed. These power modules use an innovative three-phase resonant topology that can efficiently supply full power to the thruster at an output voltage range of 200 to 700 V at an input voltage range of 80 to 160 V. Efficiencies as high as 95.9 percent were measured during an integration test with the NASA103M.XL thruster. The accuracy of the master/slave current sharing circuit and various thruster ignition techniques were evaluated.

Use of an ions thruster to dispose of type II long-lived fission products into outer space

International Nuclear Information System (INIS)

Takahashi, H.; Yu, A.

1997-01-01

To dispose of long-lived fission products (LLFPs) into outer space, an ions thruster can be used instead of a static accelerator. The specifications of the ions thrusters which are presently studies for space propulsion are presented, and their usability discussed. Using of a rocket with an ions thruster for disposing of the LLFPs directly into the sun required a larger amount of energy than does the use of an accelerator

Development of a 30-cm ion thruster thermal-vacuum power processor

Science.gov (United States)

Herron, B. G.

1976-01-01

The 30-cm Hg electron-bombardment ion thruster presently under development has reached engineering model status and is generally accepted as the prime propulsion thruster module to be used on the earliest solar electric propulsion missions. This paper presents the results of a related program to develop a transistorized 3-kW Thermal-Vacuum Breadboard (TVBB) Power Processor for this thruster. Emphasized in the paper are the implemented electrical and mechanical designs as well as the resultant system performance achieved over a range of test conditions. In addition, design modifications affording improved performance are identified and discussed.

Enabling Ring-Cusp Ion Thruster Technology for NASA Missions

Data.gov (United States)

National Aeronautics and Space Administration — ESA is flying T6 Kaufman ion thrusters on the BepiColombo Mission to Mercury in 2018. They are planning to develop a longer life, higher performing, 30-cm ring-cusp...

The physics, performance and predictions of the PEGASES ion-ion thruster

Science.gov (United States)

Aanesland, Ane

2014-10-01

Electric propulsion (EP) is now used systematically in space applications (due to the fuel and lifetime economy) to the extent that EP is now recognized as the next generation space technology. The uses of EP systems have though been limited to attitude control of GEO-stationary satellites and scientific missions. Now, the community envisages the use of EP for a variety of other applications as well; such as orbit transfer maneuvers, satellites in low altitudes, space debris removal, cube-sat control, challenging scientific missions close to and far from earth etc. For this we need a platform of EP systems providing much more variety in performance than what classical Hall and Gridded thrusters can provide alone. PEGASES is a gridded thruster that can be an alternative for some new applications in space, in particular for space debris removal. Unlike classical ion thrusters, here positive and negative ions are alternately accelerated to produce thrust. In this presentation we will look at the fundamental aspects of PEGASES. The emphasis will be put on our current understanding, obtained via analytical models, PIC simulations and experimental measurements, of the alternate extraction and acceleration process. We show that at low grid bias frequencies (10 s of kHz), the system can be described as a sequence of negative and positive ions accelerated as packets within a classical DC mode. Here secondary electrons created in the downstream chamber play an important role in the beam space charge compensation. At higher frequencies (100 s of kHz) the transit time of the ions in the grid gap becomes comparable to the bias period, leading to an ``AC acceleration mode.'' Here the beam is fully space charge compensated and the ion energy and current are functions of the applied frequency and waveform. A generalization of the Child-Langmuir space charge limited law is developed for pulsed voltages and allows evaluating the optimal parameter space and performance of PEGASES

Magnetic Field Effects on the Plume of a Diverging Cusped-Field Thruster

KAUST Repository

Matlock, Taylor

2010-07-25

The Diverging Cusped-Field Thruster (DCFT) uses three permanent ring magnets of alternating polarity to create a unique magnetic topology intended to reduce plasma losses to the discharge chamber surfaces. The magnetic field strength within the DCFT discharge chamber (up to 4 kG on axis) is much higher than in thrusters of similar geometry, which is believed to be a driving factor in the high measured anode efficiencies. The field strength in the near plume region is large as well, which may bear on the high beam divergences measured, with peaks in ion current found at angles of around 30-35 from the thruster axis. Characterization of the DCFT has heretofore involved only one magnetic topology. It is then the purpose of this study to investigate changes to the near-field plume caused by altering the shape and strength of the magnetic field. A thick magnetic collar, encircling the thruster body, is used to lower the field strength outside of the discharge chamber and thus lessen any effects caused by the external field. Changes in the thruster plume with field topology are monitored by the use of normal Langmuir and emissive probes interrogating the near-field plasma. Results are related to other observations that suggest a unified conceptual framework for the important near-exit region of the thruster.

Empirical electron cross-field mobility in a Hall effect thruster

International Nuclear Information System (INIS)

Garrigues, L.; Perez-Luna, J.; Lo, J.; Hagelaar, G. J. M.; Boeuf, J. P.; Mazouffre, S.

2009-01-01

Electron transport across the magnetic field in Hall effect thrusters is still an open question. Models have so far assumed 1/B 2 or 1/B scaling laws for the 'anomalous' electron mobility, adjusted to reproduce the integrated performance parameters of the thruster. We show that models based on such mobility laws predict very different ion velocity distribution functions (IVDF) than measured by laser induced fluorescence (LIF). A fixed spatial mobility profile, obtained by analysis of improved LIF measurements, leads to much better model predictions of thruster performance and IVDF than 1/B 2 or 1/B mobility laws for discharge voltages in the 500-700 V range.

Post-Test Inspection of Nasa's Evolutionary Xenon Thruster Long Duration Test Hardware: Ion Optics

Science.gov (United States)

Soulas, George C.; Shastry, Rohit

2016-01-01

A Long Duration Test (LDT) was initiated in June 2005 as a part of NASAs Evolutionary Xenon Thruster (NEXT) service life validation approach. Testing was voluntarily terminated in February 2014, with the thruster accumulating 51,184 hours of operation, processing 918 kg of xenon propellant, and delivering 35.5 MN-s of total impulse. This presentation will present the post-test inspection results to date for the thrusters ion optics.

High Input Voltage Discharge Supply for High Power Hall Thrusters Using Silicon Carbide Devices

Science.gov (United States)

Pinero, Luis R.; Scheidegger, Robert J.; Aulsio, Michael V.; Birchenough, Arthur G.

2014-01-01

A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because they were the best choice to handle the high voltage stress while delivering high efficiency and low specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also developed and will be integrated with the discharge supply into a vacuum-rated brassboard power processing unit with full flight functionality. This design could be evolved into a flight unit for future missions that requires high power electric propulsion.

3D ion velocity distribution function measurement in an electric thruster using laser induced fluorescence tomography

Science.gov (United States)

Elias, P. Q.; Jarrige, J.; Cucchetti, E.; Cannat, F.; Packan, D.

2017-09-01

Measuring the full ion velocity distribution function (IVDF) by non-intrusive techniques can improve our understanding of the ionization processes and beam dynamics at work in electric thrusters. In this paper, a Laser-Induced Fluorescence (LIF) tomographic reconstruction technique is applied to the measurement of the IVDF in the plume of a miniature Hall effect thruster. A setup is developed to move the laser axis along two rotation axes around the measurement volume. The fluorescence spectra taken from different viewing angles are combined using a tomographic reconstruction algorithm to build the complete 3D (in phase space) time-averaged distribution function. For the first time, this technique is used in the plume of a miniature Hall effect thruster to measure the full distribution function of the xenon ions. Two examples of reconstructions are provided, in front of the thruster nose-cone and in front of the anode channel. The reconstruction reveals the features of the ion beam, in particular on the thruster axis where a toroidal distribution function is observed. These findings are consistent with the thruster shape and operation. This technique, which can be used with other LIF schemes, could be helpful in revealing the details of the ion production regions and the beam dynamics. Using a more powerful laser source, the current implementation of the technique could be improved to reduce the measurement time and also to reconstruct the temporal evolution of the distribution function.

Spatiotemporal study of gas heating mechanisms in a radio-frequency electrothermal plasma micro-thruster

Directory of Open Access Journals (Sweden)

Amelia eGreig

2015-10-01

Full Text Available A spatiotemporal study of neutral gas temperature during the first 100 s of operation for a radio-frequency electrothermal plasma micro-thruster operating on nitrogen at 60 W and 1.5 Torr is performed to identify the heating mechanisms involved. Neutral gas temperature is estimated from rovibrational band fitting of the nitrogen second positive system. A set of baffles are used to restrict the optical image and separate the heating mechanisms occurring in the central bulk discharge region and near the thruster walls.For each spatial region there are three distinct gas heating mechanisms being fast heating from ion-neutral collisions with timescales of tens of milliseconds, intermediate heating with timescales of 10 s from ion bombardment on the inner thruster tube surface creating wall heating, and slow heating with timescales of 100 s from gradual warming of the entire thruster housing. The results are discussed in relation to optimising the thermal properties of future thruster designs.

50 KW Class Krypton Hall Thruster Performance

Science.gov (United States)

Jacobson, David T.; Manzella, David H.

2003-01-01

The performance of a 50-kilowatt-class Hall thruster designed for operation on xenon propellant was measured using kryton propellant. The thruster was operated at discharge power levels ranging from 6.4 to 72.5 kilowatts. The device produced thrust ranging from 0.3 to 2.5 newtons. The thruster was operated at discharge voltages between 250 and 1000 volts. At the highest anode mass flow rate and discharge voltage and assuming a 100 percent singly charged condition, the discharge specific impulse approached the theoretical value. Discharge specific impulse of 4500 seconds was demonstrated at a discharge voltage of 1000 volts. The peak discharge efficiency was 64 percent at 650 volts.

Two-Dimensional Modelling of the Hall Thruster Discharge: Final Report

Science.gov (United States)

2007-09-10

ion energy flux to wall, qWi, and electron energy flux to wall, qWe for Vd= 300 V, 600 V and 750 V. All variables are evaluated at the outer wall (r... qWe for Vd= 300 V, 600 V and 750 V. All variables are evaluated at the outer wall (r=0.05m). The vertical dashed line represents the thruster exit

Three Dimensional Simulation of Ion Thruster Plume-Spacecraft Interaction Based on a Graphic Processor Unit

International Nuclear Information System (INIS)

Ren Junxue; Xie Kan; Qiu Qian; Tang Haibin; Li Juan; Tian Huabing

2013-01-01

Based on the three-dimensional particle-in-cell (PIC) method and Compute Unified Device Architecture (CUDA), a parallel particle simulation code combined with a graphic processor unit (GPU) has been developed for the simulation of charge-exchange (CEX) xenon ions in the plume of an ion thruster. Using the proposed technique, the potential and CEX plasma distribution are calculated for the ion thruster plume surrounding the DS1 spacecraft at different thrust levels. The simulation results are in good agreement with measured CEX ion parameters reported in literature, and the GPU's results are equal to a CPU's. Compared with a single CPU Intel Core 2 E6300, 16-processor GPU NVIDIA GeForce 9400 GT indicates a speedup factor of 3.6 when the total macro particle number is 1.1×10 6 . The simulation results also reveal how the back flow CEX plasma affects the spacecraft floating potential, which indicates that the plume of the ion thruster is indeed able to alleviate the extreme negative floating potentials of spacecraft in geosynchronous orbit

Experimental Investigations from the Operation of a 2 Kw Brayton Power Conversion Unit and a Xenon Ion Thruster

Science.gov (United States)

Mason, Lee; Birchenough, Arthur; Pinero, Luis

2004-01-01

A 2 kW Brayton Power Conversion Unit (PCU) and a xenon ion thruster were integrated with a Power Management and Distribution (PMAD) system as part of a Nuclear Electric Propulsion (NEP) Testbed at NASA's Glenn Research Center. Brayton converters and ion thrusters are potential candidates for use on future high power NEP missions such as the proposed Jupiter Icy Moons Orbiter (JIMO). The use of existing lower power test hardware provided a cost-effective means to investigate the critical electrical interface between the power conversion system and ion propulsion system. The testing successfully demonstrated compatible electrical operations between the converter and the thruster, including end-to-end electric power throughput, high efficiency AC to DC conversion, and thruster recycle fault protection. The details of this demonstration are reported herein.

Hall-Effect Thruster Simulations with 2-D Electron Transport and Hydrodynamic Ions

Science.gov (United States)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard H.; Goebel, Dan M.

2009-01-01

A computational approach that has been used extensively in the last two decades for Hall thruster simulations is to solve a diffusion equation and energy conservation law for the electrons in a direction that is perpendicular to the magnetic field, and use discrete-particle methods for the heavy species. This "hybrid" approach has allowed for the capture of bulk plasma phenomena inside these thrusters within reasonable computational times. Regions of the thruster with complex magnetic field arrangements (such as those near eroded walls and magnets) and/or reduced Hall parameter (such as those near the anode and the cathode plume) challenge the validity of the quasi-one-dimensional assumption for the electrons. This paper reports on the development of a computer code that solves numerically the 2-D axisymmetric vector form of Ohm's law, with no assumptions regarding the rate of electron transport in the parallel and perpendicular directions. The numerical challenges related to the large disparity of the transport coefficients in the two directions are met by solving the equations in a computational mesh that is aligned with the magnetic field. The fully-2D approach allows for a large physical domain that extends more than five times the thruster channel length in the axial direction, and encompasses the cathode boundary. Ions are treated as an isothermal, cold (relative to the electrons) fluid, accounting for charge-exchange and multiple-ionization collisions in the momentum equations. A first series of simulations of two Hall thrusters, namely the BPT-4000 and a 6-kW laboratory thruster, quantifies the significance of ion diffusion in the anode region and the importance of the extended physical domain on studies related to the impact of the transport coefficients on the electron flow field.

Reduced power processor requirements for the 30-cm diameter HG ion thruster

Science.gov (United States)

Rawlin, V. K.

1979-01-01

The characteristics of power processors strongly impact the overall performance and cost of electric propulsion systems. A program was initiated to evaluate simplifications of the thruster-power processor interface requirements. The power processor requirements are mission dependent with major differences arising for those missions which require a nearly constant thruster operating point (typical of geocentric and some inbound planetary missions) and those requiring operation over a large range of input power (such as outbound planetary missions). This paper describes the results of tests which have indicated that as many as seven of the twelve power supplies may be eliminated from the present Functional Model Power Processor used with 30-cm diameter Hg ion thrusters.

Pickup ion processes associated with spacecraft thrusters: Implications for solar probe plus

Energy Technology Data Exchange (ETDEWEB)

Clemens, Adam, E-mail: a.j.clemens@qmul.ac.uk; Burgess, David [School of Physics and Astronomy, Queen Mary University of London, London (United Kingdom)

2016-03-15

Chemical thrusters are widely used in spacecraft for attitude control and orbital manoeuvres. They create an exhaust plume of neutral gas which produces ions via photoionization and charge exchange. Measurements of local plasma properties will be affected by perturbations caused by the coupling between the newborn ions and the plasma. A model of neutral expansion has been used in conjunction with a fully three-dimensional hybrid code to study the evolution and ionization over time of the neutral cloud produced by the firing of a mono-propellant hydrazine thruster as well as the interactions of the resulting ion cloud with the ambient solar wind. Results are presented which show that the plasma in the region near to the spacecraft will be perturbed for an extended period of time with the formation of an interaction region around the spacecraft, a moderate amplitude density bow wave bounding the interaction region and evidence of an instability at the forefront of the interaction region which causes clumps of ions to be ejected from the main ion cloud quasi-periodically.

Computational simulation of coupled nonequilibrium discharge and compressible flow phenomena in a microplasma thruster

International Nuclear Information System (INIS)

Deconinck, Thomas; Mahadevan, Shankar; Raja, Laxminarayan L.

2009-01-01

The microplasma thruster (MPT) concept is a simple extension of a cold gas micronozzle propulsion device, where a direct-current microdischarge is used to preheat the gas stream to improve the specific impulse of the device. Here we study a prototypical MPT device using a detailed, self-consistently coupled plasma and flow computational model. The model describes the microdischarge power deposition, plasma dynamics, gas-phase chemical kinetics, coupling of the plasma phenomena with high-speed flow, and overall propulsion system performance. Compared to a cold gas micronozzle, a significant increase in specific impulse is obtained from the power deposition in the diverging section of the MPT nozzle. For a discharge voltage of 750 V, a power input of 650 mW, and an argon mass flow rate of 5 SCCM (SCCM denotes cubic centimeter per minute at STP), the specific impulse of the device is increased by a factor of ∼1.5 to about 74 s. The microdischarge remains mostly confined inside the micronozzle and operates in an abnormal glow discharge regime. Gas heating, primarily due to ion Joule heating, is found to have a strong influence on the overall discharge behavior. The study provides a validation of the MPT concept as a simple and effective approach to improve the performance of micronozzle cold gas propulsion devices.

Experimental Investigation of the Near-Wall Region in the NASA HiVHAc EDU2 Hall Thruster

Science.gov (United States)

Shastry, Rohit; Kamhawi, Hani; Huang, Wensheng; Haag, Thomas W.

2015-01-01

The HiVHAc propulsion system is currently being developed to support Discovery-class NASA science missions. Presently, the thruster meets the required operational lifetime by utilizing a novel discharge channel replacement mechanism. As a risk reduction activity, an alternative approach is being investigated that modifies the existing magnetic circuit to shift the ion acceleration zone further downstream such that the magnetic components are not exposed to direct ion impingement during the thruster's lifetime while maintaining adequate thruster performance and stability. To measure the change in plasma properties between the original magnetic circuit configuration and the modified, "advanced" configuration, six Langmuir probes were flush-mounted within each channel wall near the thruster exit plane. Plasma potential and electron temperature were measured for both configurations across a wide range of discharge voltages and powers. Measurements indicate that the upstream edge of the acceleration zone shifted downstream by as much as 0.104 channel lengths, depending on operating condition. The upstream edge of the acceleration zone also appears to be more insensitive to operating condition in the advanced configuration, remaining between 0.136 and 0.178 channel lengths upstream of the thruster exit plane. Facility effects studies performed on the original configuration indicate that the plasma and acceleration zone recede further upstream into the channel with increasing facility pressure. These results will be used to inform further modifications to the magnetic circuit that will provide maximum protection of the magnetic components without significant changes to thruster performance and stability.

Rarefied gas electro jet (RGEJ) micro-thruster for space propulsion

International Nuclear Information System (INIS)

Blanco, Ariel; Roy, Subrata

2017-01-01

This article numerically investigates a micro-thruster for small satellites which utilizes plasma actuators to heat and accelerate the flow in a micro-channel with rarefied gas in the slip flow regime. The inlet plenum condition is considered at 1 Torr with flow discharging to near vacuum conditions (<0.05 Torr). The Knudsen numbers at the inlet and exit planes are ∼0.01 and ∼0.1, respectively. Although several studies have been performed in micro-hallow cathode discharges at constant pressure, to our knowledge, an integrated study of the glow discharge physics and resulting fluid flow of a plasma thruster under these low pressure and low Knudsen number conditions is yet to be reported. Numerical simulations of the charge distribution due to gas ionization processes and the resulting rarefied gas flow are performed using an in-house code. The mass flow rate, thrust, specific impulse, power consumption and the thrust effectiveness of the thruster are predicted based on these results. The ionized gas is modelled using local mean energy approximation. An electrically induced body force and a thermal heating source are calculated based on the space separated charge distribution and the ion Joule heating, respectively. The rarefied gas flow with these electric force and heating source is modelled using density-based compressible flow equations with slip flow boundary conditions. The results show that a significant improvement of specific impulse can be achieved over highly optimized cold gas thrusters using the same propellant. (paper)

Effects of Enhanced Eathode Electron Emission on Hall Thruster Operation

International Nuclear Information System (INIS)

Raitses, Y.; Smirnov, A.; Fisch, N.J.

2009-01-01

Interesting discharge phenomena are observed that have to do with the interaction between the magnetized Hall thruster plasma and the neutralizing cathode. The steadystate parameters of a highly ionized thruster discharge are strongly influenced by the electron supply from the cathode. The enhancement of the cathode electron emission above its self-sustained level affects the discharge current and leads to a dramatic reduction of the plasma divergence and a suppression of large amplitude, low frequency discharge current oscillations usually related to an ionization instability. These effects correlate strongly with the reduction of the voltage drop in the region with the fringing magnetic field between the thruster channel and the cathode. The measured changes of the plasma properties suggest that the electron emission affects the electron cross-field transport in the thruster discharge. These trends are generalized for Hall thrusters of various configurations.

Low frequency azimuthal stability of the ionization region of the Hall thruster discharge. II. Global analysis

International Nuclear Information System (INIS)

Escobar, D.; Ahedo, E.

2015-01-01

The linear stability of the Hall thruster discharge is analysed against axial-azimuthal perturbations in the low frequency range using a time-dependent 2D code of the discharge. This azimuthal stability analysis is spatially global, as opposed to the more common local stability analyses, already afforded previously (D. Escobar and E. Ahedo, Phys. Plasmas 21(4), 043505 (2014)). The study covers both axial and axial-azimuthal oscillations, known as breathing mode and spoke, respectively. The influence on the spoke instability of different operation parameters such as discharge voltage, mass flow, and thruster size is assessed by means of different parametric variations and compared against experimental results. Additionally, simplified models are used to unveil and characterize the mechanisms driving the spoke. The results indicate that the spoke is linked to azimuthal oscillations of the ionization process and to the Bohm condition in the transition to the anode sheath. Finally, results obtained from local and global stability analyses are compared in order to explain the discrepancies between both methods

Plasma Structure and Behavior of Miniature Ring-Cusp Discharges

Science.gov (United States)

Mao, Hann-Shin

Miniature ring-cusp ion thrusters provide a unique blend of high efficiencies and millinewton level thrust for future spacecraft. These thrusters are attractive as a primary propulsion for small satellites that require a high delta V, and as a secondary propulsion for larger spacecraft that require precision formation flying, disturbance rejection, or attitude control. To ensure desirable performance throughout the life of such missions, an advancement in the understanding of the plasma structure and behavior of miniature ring-cusp discharges is required. A research model was fabricated to provide a simplified experimental test bed for the analysis of the plasma discharge chamber of a miniature ion thruster. The plasma source allowed for spatially resolved measurements with a Langmuir probe along a meridian plane. Probe measurements yielded plasma density, electron temperature, and plasma potential data. The magnetic field strength was varied along with the discharge current to determine the plasma behavior under various conditions. The structure of the plasma properties were found to be independent of the discharge power under the proper scaling. It was concluded that weaker magnetic fields can improve the overall performance for ion thruster operation. To further analyze the experimental measurements, a framework was developed based on the magnetic field. A flux aligned coordinate system was developed to decouple the perpendicular and parallel plasma motion with respect to the magnetic field. This was done using the stream function and magnetic scalar potential. Magnetic formulae provided intuition on the field profiles dependence on magnet dimensions. The flux aligned coordinate system showed that the plasma was isopycnic along constant stream function values. This was used to develop an empirical relation suitable for estimating the spatial behavior and to determine the plasma volume and loss areas. The plasma geometry estimates were applied to a control volume

Magnetically filtered Faraday probe for measuring the ion current density profile of a Hall thruster

International Nuclear Information System (INIS)

Rovey, Joshua L.; Walker, Mitchell L.R.; Gallimore, Alec D.; Peterson, Peter Y.

2006-01-01

The ability of a magnetically filtered Faraday probe (MFFP) to obtain the ion current density profile of a Hall thruster is investigated. The MFFP is designed to eliminate the collection of low-energy, charge-exchange (CEX) ions by using a variable magnetic field as an ion filter. In this study, a MFFP, Faraday probe with a reduced acceptance angle (BFP), and nude Faraday probe are used to measure the ion current density profile of a 5 kW Hall thruster operating over the range of 300-500 V and 5-10 mg/s. The probes are evaluated on a xenon propellant Hall thruster in the University of Michigan Large Vacuum Test Facility at operating pressures within the range of 4.4x10 -4 Pa Xe (3.3x10 -6 Torr Xe) to 1.1x10 -3 Pa Xe (8.4x10 -6 Torr Xe) in order to study the ability of the Faraday probe designs to filter out CEX ions. Detailed examination of the results shows that the nude probe measures a greater ion current density profile than both the MFFP and BFP over the range of angular positions investigated for each operating condition. The differences between the current density profiles obtained by each probe are attributed to the ion filtering systems employed. Analysis of the results shows that the MFFP, operating at a +5 A solenoid current, provides the best agreement with flight-test data and across operating pressures

Comparison of hollow cathode discharge plasma configurations

International Nuclear Information System (INIS)

Farnell, Casey C; Farnell, Cody C; Williams, John D

2011-01-01

Hollow cathodes used in plasma contactor and electric propulsion devices provide electrons for sustaining plasma discharges and enabling plasma bridge neutralization. Life tests show erosion on hollow cathodes exposed to the plasma environment produced in the region downstream of these devices. To explain the observed erosion, plasma flow field measurements are presented for hollow cathode generated plasmas using both directly immersed probes and remotely located plasma diagnostics. Measurements on two cathode discharge configurations are presented: (1) an open, no magnetic field configuration and (2) a setup simulating the discharge chamber environment of an ion thruster. In the open cathode configuration, large amplitude plasma potential oscillations, ranging from 20 to 85 V within a 34 V discharge, were observed using a fast response emissive probe. These oscillations were observed over a dc potential profile that included a well-defined potential hill structure. A remotely located electrostatic analyzer (ESA) was used to measure the energy of ions produced within the plasma, and energies were detected that met, and in some cases exceeded, the peak oscillatory plasma potentials detected by the emissive probe. In the ion thruster discharge chamber configuration, plasma potentials from the emissive probe again agreed with ion energies recorded by the remotely located ESA; however, much lower ion energies were detected compared with the open configuration. A simplified ion-transit model that uses temporal and spatial plasma property measurements is presented and used to predict far-field plasma streaming properties. Comparisons between the model and remote measurements are presented.

NASA's Evolutionary Xenon Thruster (NEXT) Project Qualification Propellant Throughput Milestone: Performance, Erosion, and Thruster Service Life Prediction After 450 kg

Science.gov (United States)

Herman, Daniel A.

2010-01-01

The NASA s Evolutionary Xenon Thruster (NEXT) program is tasked with significantly improving and extending the capabilities of current state-of-the-art NSTAR thruster. The service life capability of the NEXT ion thruster is being assessed by thruster wear test and life-modeling of critical thruster components, such as the ion optics and cathodes. The NEXT Long-Duration Test (LDT) was initiated to validate and qualify the NEXT thruster propellant throughput capability. The NEXT thruster completed the primary goal of the LDT; namely to demonstrate the project qualification throughput of 450 kg by the end of calendar year 2009. The NEXT LDT has demonstrated 28,500 hr of operation and processed 466 kg of xenon throughput--more than double the throughput demonstrated by the NSTAR flight-spare. Thruster performance changes have been consistent with a priori predictions. Thruster erosion has been minimal and consistent with the thruster service life assessment, which predicts the first failure mode at greater than 750 kg throughput. The life-limiting failure mode for NEXT is predicted to be loss of structural integrity of the accelerator grid due to erosion by charge-exchange ions.

Low power arcjet thruster pulse ignition

Science.gov (United States)

Sarmiento, Charles J.; Gruber, Robert P.

1987-01-01

An investigation of the pulse ignition characteristics of a 1 kW class arcjet using an inductive energy storage pulse generator with a pulse width modulated power converter identified several thruster and pulse generator parameters that influence breakdown voltage including pulse generator rate of voltage rise. This work was conducted with an arcjet tested on hydrogen-nitrogen gas mixtures to simulate fully decomposed hydrazine. Over all ranges of thruster and pulser parameters investigated, the mean breakdown voltages varied from 1.4 to 2.7 kV. Ignition tests at elevated thruster temperatures under certain conditions revealed occasional breakdowns to thruster voltages higher than the power converter output voltage. These post breakdown discharges sometimes failed to transition to the lower voltage arc discharge mode and the thruster would not ignite. Under the same conditions, a transition to the arc mode would occur for a subsequent pulse and the thruster would ignite. An automated 11 600 cycle starting and transition to steady state test demonstrated ignition on the first pulse and required application of a second pulse only two times to initiate breakdown.

Experimental Investigation from the Operation of a 2 kW Brayton Power Conversion Unit and a Xenon Ion Thruster

Science.gov (United States)

Hervol, David; Mason, Lee; Birchenough, Art; Pinero, Luis

2004-01-01

A 2kW Brayton Power Conversion Unit (PCU) and a xenon ion thruster were integrated with a Power Management and Distribution (PMAD) system as part of a Nuclear Electric Propulsion (NEP) Testbed at NASA's Glenn Research Center. Brayton Converters and ion thrusters are potential candidates for use on future high power NEP mission such as the proposed Jupiter Icy Moons Orbiter (JIMO). The use of a existing lower power test hardware provided a cost effective means to investigate the critical electrical interface between the power conversion system and the propulsion system. The testing successfully demonstrated compatible electrical operations between the converter and the thruster, including end-to-end electric power throughput, high efficiency AC to DC conversion, and thruster recycle fault protection. The details of this demonstration are reported herein.

Ion velocities in a micro-cathode arc thruster

International Nuclear Information System (INIS)

Zhuang Taisen; Shashurin, Alexey; Keidar, Michael; Beilis, Isak

2012-01-01

Ion velocities in the plasma jet generated by the micro-cathode arc thruster are studied by means of time-of-flight method using enhanced ion detection system (EIDS). The EIDS triggers perturbations (spikes) on arc current waveform, and the larger current in the spike generates denser plasma bunches propagating along with the mainstream plasma. The EIDS utilizes double electrostatic probes rather than single probes. The average Ti ion velocity is measured to be around 2×10 4 m/s without a magnetic field. It was found that the application of a magnetic field does not change ion velocities in the interelectrode region while leads to ion acceleration in the free expanding plasma plume by a factor of about 2. Ion velocities of about 3.5×10 4 m/s were detected for the magnetic field of about 300 mT at distance of about 100–200 mm from the cathode. It is proposed that plasma is accelerated due to Lorentz force. The average thrust is calculated using the ion velocity measurements and the cathode mass consumption rate, and its increase with the magnetic field is demonstrated.

Development of an Ion Thruster and Power Processor for New Millennium's Deep Space 1 Mission

Science.gov (United States)

Sovey, James S.; Hamley, John A.; Haag, Thomas W.; Patterson, Michael J.; Pencil, Eric J.; Peterson, Todd T.; Pinero, Luis R.; Power, John L.; Rawlin, Vincent K.; Sarmiento, Charles J.;

Simulation of Main Plasma Parameters of a Cylindrical Asymmetric Capacitively Coupled Plasma Micro-Thruster using Computational Fluid Dynamics

Directory of Open Access Journals (Sweden)

Amelia eGreig

2015-01-01

Full Text Available Computational fluid dynamics (CFD simulations of a radio-frequency (13.56 MHz electro-thermal capacitively coupled plasma (CCP micro-thruster have been performed using the commercial CFD-ACE+ package. Standard operating conditions of a 10 W, 1.5 Torr argon discharge were used to compare with previously obtained experimental results for validation. Results show that the driving force behind plasma production within the thruster is ion-induced secondary electrons ejected from the surface of the discharge tube, accelerated through the sheath to electron temperatures up to 33.5 eV. The secondary electron coefficient was varied to determine the effect on the discharge, with results showing that full breakdown of the discharge did not occur for coefficients coefficients less than or equal to 0.01.

Improvement of the low frequency oscillation model for Hall thrusters

Energy Technology Data Exchange (ETDEWEB)

Wang, Chunsheng, E-mail: wangcs@hit.edu.cn; Wang, Huashan [Yanshan University, College of Vehicles and Energy, Qinhuangdao 066004, Hebei (China)

2016-08-15

The low frequency oscillation of the discharge current in Hall thrusters is a major aspect of these devices that requires further study. While the existing model captures the ionization mechanism of the low frequency oscillation, it unfortunately fails to express the dynamic characteristics of the ion acceleration. The analysis in this paper shows this is because of the simplification of the electron equation, which affects both the electric field distribution and the ion acceleration process. Additionally, the electron density equation is revised and a new model that is based on the physical properties of ion movement is proposed.

ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS IN HALL EFFECT THRUSTERS

Science.gov (United States)

2017-06-30

NUMBER (Include area code) 30 June 2017 Briefing Charts 26 May 2017 - 30 June 2017 ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS ...Robert Martin N/A ION ACOUSTIC TURBULENCE, ANOMALOUS TRANSPORT, AND SYSTEM DYNAMICS IN HALL EFFECT THRUSTERS Robert Martin1, Jonathan Tran2 1AIR FORCE...Approved for Public Release; Distribution is Unlimited. PA# 17394 1 / 13 OUTLINE 1 INTRODUCTION 2 TRANSPORT 3 DYNAMIC SYSTEM 4 SUMMARY AND CONCLUSION

Rarefied gas electro jet (RGEJ) micro-thruster for space propulsion

Science.gov (United States)

Blanco, Ariel; Roy, Subrata

2017-11-01

This article numerically investigates a micro-thruster for small satellites which utilizes plasma actuators to heat and accelerate the flow in a micro-channel with rarefied gas in the slip flow regime. The inlet plenum condition is considered at 1 Torr with flow discharging to near vacuum conditions (consumption and the thrust effectiveness of the thruster are predicted based on these results. The ionized gas is modelled using local mean energy approximation. An electrically induced body force and a thermal heating source are calculated based on the space separated charge distribution and the ion Joule heating, respectively. The rarefied gas flow with these electric force and heating source is modelled using density-based compressible flow equations with slip flow boundary conditions. The results show that a significant improvement of specific impulse can be achieved over highly optimized cold gas thrusters using the same propellant.

Anode sheath in Hall thrusters

International Nuclear Information System (INIS)

Dorf, L.; Semenov, V.; Raitses, Y.

2003-01-01

A set of hydrodynamic equations is used to describe quasineutral plasma in ionization and acceleration regions of a Hall thruster. The electron distribution function and Poisson equation are invoked for description of a near-anode region. Numerical solutions suggest that steady-state operation of a Hall thruster can be achieved at different anode sheath regimes. It is shown that the anode sheath depends on the thruster operating conditions, namely the discharge voltage and the mass flow rate

Internal plasma potential measurements of a Hall thruster using xenon and krypton propellant

International Nuclear Information System (INIS)

Linnell, Jesse A.; Gallimore, Alec D.

2006-01-01

For krypton to become a realistic option for Hall thruster operation, it is necessary to understand the performance gap between xenon and krypton and what can be done to reduce it. A floating emissive probe is used with the Plasmadynamics and Electric Propulsion Laboratory's High-speed Axial Reciprocating Probe system to map the internal plasma potential structure of the NASA-173Mv1 Hall thruster [R. R. Hofer, R. S. Jankovsky, and A. D. Gallimore, J. Propulsion Power 22, 721 (2006); and ibid.22, 732 (2006)] using xenon and krypton propellant. Measurements are taken for both propellants at discharge voltages of 500 and 600 V. Electron temperatures and electric fields are also reported. The acceleration zone and equipotential lines are found to be strongly linked to the magnetic-field lines. The electrostatic plasma lens of the NASA-173Mv1 Hall thruster strongly focuses the xenon ions toward the center of the discharge channel, whereas the krypton ions are defocused. Krypton is also found to have a longer acceleration zone than the xenon cases. These results explain the large beam divergence observed with krypton operation. Krypton and xenon have similar maximum electron temperatures and similar lengths of the high electron temperature zone, although the high electron temperature zone is located farther downstream in the krypton case

Performance, Facility Pressure Effects, and Stability Characterization Tests of NASA's Hall Effect Rocket with Magnetic Shielding Thruster

Science.gov (United States)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Peterson, Peter; Hofer, Richard; Mikellides, Ioannis

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster 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 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 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 was mostly invariant to changes in the facility background pressure for vacuum chamber pressure below 110-5 Torr-Xe (for thruster flow rate above 8 mgs). Power spectral density analysis of the discharge current waveform showed that increasing the vacuum chamber background pressure resulted in a higher discharge current dominant frequency. Finally the IVB maps of the TDU-1

One-dimensional hybrid-direct kinetic simulation of the discharge plasma in a Hall thruster

International Nuclear Information System (INIS)

Hara, Kentaro; Boyd, Iain D.; Kolobov, Vladimir I.

2012-01-01

In order to model the non-equilibrium plasma within the discharge region of a Hall thruster, the velocity distribution functions (VDFs) must be obtained accurately. A direct kinetic (DK) simulation method that directly solves the plasma Boltzmann equation can achieve better resolution of VDFs in comparison to particle simulations, such as the particle-in-cell (PIC) method that inherently include statistical noise. In this paper, a one-dimensional hybrid-DK simulation, which uses a DK simulation for heavy species and a fluid model for electrons, is developed and compared to a hybrid-PIC simulation. Time-averaged results obtained from the hybrid-DK simulation are in good agreement with hybrid-PIC results and experimental data. It is shown from a comparison of using a kinetic simulation and solving the continuity equation that modeling of the neutral atoms plays an important role for simulations of the Hall thruster discharge plasma. In addition, low and high frequency plasma oscillations are observed. Although the kinetic nature of electrons is not resolved due to the use of a fluid model, the hybrid-DK model provides spatially and temporally well-resolved plasma properties and an improved resolution of VDFs for heavy species with less statistical noise in comparison to the hybrid-PIC method.

Experiments on Ion-Ion Plasmas From Discharges

Science.gov (United States)

Leonhardt, Darrin; Walton, Scott; Blackwell, David; Murphy, Donald; Fernsler, Richard; Meger, Robert

2001-10-01

Use of both positive and negative ions in plasma processing of materials has been shown to be advantageous[1] in terms of better feature evolution and control. In this presentation, experimental results are given to complement recent theoretical work[2] at NRL on the formation and decay of pulsed ion-ion plasmas in electron beam generated discharges. Temporally resolved Langmuir probe and mass spectrometry are used to investigate electron beam generated discharges during the beam on (active) and off (afterglow) phases in a variety of gas mixtures. Because electron-beam generated discharges inherently[3] have low electron temperatures (<0.5eV in molecular gases), negative ion characteristics are seen in the active as well as afterglow phases since electron detachment increases with low electron temperatures. Analysis of temporally resolved plasma characteristics deduced from these measurements will be presented for pure O_2, N2 and Ar and their mixtures with SF_6. Oxygen discharges show no noticeable negative ion contribution during the active or afterglow phase, presumably due to the higher energy electron attachment threshold, which is well above any electron temperature. In contrast, SF6 discharges demonstrate ion-ion plasma characteristics in the active glow and are completely ion-ion in the afterglow. Comparison between these discharges with published cross sections and production mechanisms will also be presented. [1] T.H. Ahn, K. Nakamura & H. Sugai, Plasma Sources Sci. Technol., 5, 139 (1996); T. Shibyama, H. Shindo & Y. Horiike, Plasma Sources Sci. Technol., 5, 254 (1996). [2] See presentation by R. F. Fernsler, at this conference. [3] D. Leonhardt, et al., 53rd Annual GEC, Houston, TX.

Ion source of discharge type

Energy Technology Data Exchange (ETDEWEB)

Enchevich, I.B. [TRIUMF, Cyclotron Div., Vancouver, British Columbia (Canada); Korenev, S.A. [JINR, Hihg Energy Physics Lab., Dubna, Moscow (Russian Federation)

1992-07-01

A new scheme of ion source based on a dielectric surface sliding discharge is described. The conditions to form this type of discharge are analyzed and experimental results are shown. The main parameters of this ion source are: accelerating voltage U = 1/20kV; continuous extracted ion beam; current density j = 0.01/0.5 A/cm{sup 2}; ions of Cl, F, C, H; residual gas pressure P = 10{sup -6} Torr. A magnetic system is used to separate the different types of ions. The dielectric material in the discharge circuit (anode plasma emitter) defines the type of ions. The emission characteristics of plasma emitter and the discharge parameters are presented. The ion current yield satisfies the Child-Langmuir law. (author)

Ion source of discharge type

International Nuclear Information System (INIS)

Enchevich, I.B.; Korenev, S.A.

1992-07-01

A new scheme of ion source based on a dielectric surface sliding discharge is described. The conditions to form this type of discharge are analyzed and experimental results are shown. The main parameters of this ion source are: accelerating voltage U = 1/20kV; continuous extracted ion beam; current density j = 0.01/0.5 A/cm 2 ; ions of Cl, F, C, H; residual gas pressure P = 10 -6 Torr. A magnetic system is used to separate the different types of ions. The dielectric material in the discharge circuit (anode plasma emitter) defines the type of ions. The emission characteristics of plasma emitter and the discharge parameters are presented. The ion current yield satisfies the Child-Langmuir law. (author)

Optical Diagnostic Characterization of High-Power Hall Thruster Wear and Operation

Science.gov (United States)

Williams, George J., Jr.; Soulas, George C.; Kamhawi, Hani

2012-01-01

Optical emission spectroscopy is employed to correlate BN insulator erosion with high-power Hall thruster operation. Specifically, actinometry leveraging excited xenon states is used to normalize the emission spectra of ground state boron as a function of thruster operating condition. Trends in the strength of the boron signal are correlated with thruster power, discharge voltage, and discharge current. In addition, the technique is demonstrated on metallic coupons embedded in the walls of the HiVHAc EM thruster. The OES technique captured the overall trend in the erosion of the coupons which boosts credibility in the method since there are no data to which to calibrate the erosion rates of high-power Hall thrusters. The boron signals are shown to trend linearly with discharge voltage for a fixed discharge current as expected. However, the boron signals of the higher-power NASA 300M and NASA 457Mv2 trend with discharge current and show an unexpectedly weak to inverse dependence on discharge voltage. Electron temperatures measured optically in the near-field plume of the thruster agree well with Langmuir probe data. However, the optical technique used to determine Te showed unacceptable sensitivity to the emission intensities. Near-field, single-frequency imaging of the xenon neutrals is also presented as a function of operating condition for the NASA 457 Mv2.

Discharge characteristics of an ablative pulsed plasma thruster with non-volatile liquid propellant

Science.gov (United States)

Ling, William Yeong Liang; Schönherr, Tony; Koizumi, Hiroyuki

2017-07-01

Pulsed plasma thrusters (PPTs) are a form of electric spacecraft propulsion. They have an extremely simple structure and are highly suitable for nano/micro-spacecraft with weights in the kilogram range. Such small spacecraft have recently experienced increased growth but still lack suitable efficient propulsion systems. PPTs operate in a pulsed mode (one discharge = one shot) and typically use solid polytetrafluoroethylene (PTFE) as a propellant. However, new non-volatile liquids in the perfluoropolyether (PFPE) family have recently been found to be promising alternatives. A recent study presented results on the physical characteristics of PFPE vs. PTFE, showing that PFPE is superior in terms of physical characteristics such as its resistance to carbon deposition. This letter will examine the electrical discharge characteristics of PFPE vs. PTFE. The results demonstrate that PFPE has excellent shot-to-shot repeatability and a lower discharge resistance when compared with PTFE. Taken together with its physical characteristics, PFPE appears to be a strong contender to PTFE as a PPT propellant.

Comparisons in Performance of Electromagnet and Permanent-Magnet Cylindrical Hall-Effect Thrusters

Science.gov (United States)

Polzin, K. A.; Raitses, Y.; Gayoso, J. C.; Fisch, N. J.

2010-01-01

Three different low-power cylindrical Hall thrusters, which more readily lend themselves to miniaturization and low-power operation than a conventional (annular) Hall thruster, are compared to evaluate the propulsive performance of each. One thruster uses electromagnet coils to produce the magnetic field within the discharge channel while the others use permanent magnets, promising power reduction relative to the electromagnet thruster. A magnetic screen is added to the permanent magnet thruster to improve performance by keeping the magnetic field from expanding into space beyond the exit of the thruster. The combined dataset spans a power range from 50-350 W. The thrust levels over this range were 1.3-7.3 mN, with thruster efficiencies and specific impulses spanning 3.5-28.7% and 400-1940 s, respectively. The efficiency is generally higher for the permanent magnet thruster with the magnetic screen, while That thruster s specific impulse as a function of discharge voltage is comparable to the electromagnet thruster.

A centre-triggered magnesium fuelled cathodic arc thruster uses sublimation to deliver a record high specific impulse

Science.gov (United States)

Neumann, Patrick R. C.; Bilek, Marcela; McKenzie, David R.

2016-08-01

The cathodic arc is a high current, low voltage discharge that operates in vacuum and provides a stream of highly ionised plasma from a solid conducting cathode. The high ion velocities, together with the high ionisation fraction and the quasineutrality of the exhaust stream, make the cathodic arc an attractive plasma source for spacecraft propulsion applications. The specific impulse of the cathodic arc thruster is substantially increased when the emission of neutral species is reduced. Here, we demonstrate a reduction of neutral emission by exploiting sublimation in cathode spots and enhanced ionisation of the plasma in short, high-current pulses. This, combined with the enhanced directionality due to the efficient erosion profiles created by centre-triggering, substantially increases the specific impulse. We present experimentally measured specific impulses and jet power efficiencies for titanium and magnesium fuels. Our Mg fuelled source provides the highest reported specific impulse for a gridless ion thruster and is competitive with all flight rated ion thrusters. We present a model based on cathode sublimation and melting at the cathodic arc spot explaining the outstanding performance of the Mg fuelled source. A further significant advantage of an Mg-fuelled thruster is the abundance of Mg in asteroidal material and in space junk, providing an opportunity for utilising these resources in space.

Advanced electrostatic ion thruster for space propulsion

Science.gov (United States)

Masek, T. D.; Macpherson, D.; Gelon, W.; Kami, S.; Poeschel, R. L.; Ward, J. W.

1978-01-01

The suitability of the baseline 30 cm thruster for future space missions was examined. Preliminary design concepts for several advanced thrusters were developed to assess the potential practical difficulties of a new design. Useful methodologies were produced for assessing both planetary and earth orbit missions. Payload performance as a function of propulsion system technology level and cost sensitivity to propulsion system technology level are among the topics assessed. A 50 cm diameter thruster designed to operate with a beam voltage of about 2400 V is suggested to satisfy most of the requirements of future space missions.

Experimental test of 200 W Hall thruster with titanium wall

Science.gov (United States)

Ding, Yongjie; Sun, Hezhi; Peng, Wuji; Xu, Yu; Wei, Liqiu; Li, Hong; Li, Peng; Su, Hongbo; Yu, Daren

2017-05-01

We designed a 200 W Hall thruster based on the technology of pushing down a magnetic field with two permanent magnetic rings. Boron nitride (BN) is an important insulating wall material for Hall thrusters. The discharge characteristics of the designed Hall thruster were studied by replacing BN with titanium (Ti). Experimental results show that the designed Hall thruster can discharge stably for a long time under a Ti channel. Experiments were performed to determine whether the channel and cathode are electrically connected. When the channel wall and cathode are insulated, the divergence angle of the plume increases, but the performance of the Hall thruster is improved in terms of thrust, specific impulse, anode efficiency, and thrust-to-power ratio. Ti exhibits a powerful antisputtering capability, a low emanation rate of gas, and a large structural strength, making it a potential candidate wall material in the design of low-power Hall thrusters.

Low-Cost, High-Performance Hall Thruster Support System

Science.gov (United States)

Hesterman, Bryce

2015-01-01

Colorado Power Electronics (CPE) has built an innovative modular PPU for Hall thrusters, including discharge, magnet, heater and keeper supplies, and an interface module. This high-performance PPU offers resonant circuit topologies, magnetics design, modularity, and a stable and sustained operation during severe Hall effect thruster current oscillations. Laboratory testing has demonstrated discharge module efficiency of 96 percent, which is considerably higher than current state of the art.

Investigations of Probe Induced Perturbations in a Hall Thruster

International Nuclear Information System (INIS)

D. Staack; Y. Raitses; N.J. Fisch

2002-01-01

An electrostatic probe used to measure spatial plasma parameters in a Hall thruster generates perturbations of the plasma. These perturbations are examined by varying the probe material, penetration distance, residence time, and the nominal thruster conditions. The study leads us to recommendations for probe design and thruster operating conditions to reduce discharge perturbations, including metal shielding of the probe insulator and operation of the thruster at lower densities

Performance Test Results of the NASA-457M v2 Hall Thruster

Science.gov (United States)

Soulas, George C.; Haag, Thomas W.; Herman, Daniel A.; Huang, Wensheng; Kamhawi, Hani; Shastry, Rohit

2012-01-01

Performance testing of a second generation, 50 kW-class Hall thruster labeled NASA-457M v2 was conducted at the NASA Glenn Research Center. This NASA-designed thruster is an excellent candidate for a solar electric propulsion system that supports human exploration missions. Thruster discharge power was varied from 5 to 50 kW over discharge voltage and current ranges of 200 to 500 V and 15 to 100 A, respectively. Anode efficiencies varied from 0.56 to 0.71. The peak efficiency was similar to that of other state-of-the-art high power Hall thrusters, but outperformed these thrusters at lower discharge voltages. The 0.05 to 0.18 higher anode efficiencies of this thruster compared to its predecessor were primarily due to which of two stable discharge modes the thruster was operated. One stable mode was at low magnetic field strengths, which produced high anode efficiencies, and the other at high magnetic fields where its predecessor was operated. Cathode keeper voltages were always within 2.1 to 6.2 V and cathode voltages were within 13 V of tank ground during high anode efficiency operation. However, during operation at high magnetic fields, cathode-to-ground voltage magnitudes increased dramatically, exceeding 30 V, due to the high axial magnetic field strengths in the immediate vicinity of the centrally-mounted cathode. The peak thrust was 2.3 N and this occurred at a total thruster input power of 50.0 kW at a 500 V discharge voltage. The thruster demonstrated a thrust-to-power range of 76.4 mN/kW at low power to 46.1 mN/kW at full power, and a specific impulse range of 1420 to 2740 s. For a discharge voltage of 300 V, where specific impulses would be about 2000 s, thrust efficiencies varied from 0.57 to 0.63.

Study of the key factors affecting the triple grid lifetime of the LIPS-300 ion thruster

Science.gov (United States)

Mingming, SUN; Liang, WANG; Juntai, YANG; Xiaodong, WEN; Yongjie, HUANG; Meng, WANG

2018-04-01

In order to ascertain the key factors affecting the lifetime of the triple grids in the LIPS-300 ion thruster, the thermal deformation, upstream ion density and component lifetime of the grids are simulated with finite element analysis, fluid simulation and charged-particle tracing simulation methods on the basis of a 1500 h short lifetime test. The key factor affecting the lifetime of the triple grids in the LIPS-300 ion thruster is obtained and analyzed through the test results. The results show that ion sputtering erosion of the grids in 5 kW operation mode is greater than in the case of 3 kW. In 5 kW mode, the decelerator grid shows the most serious corrosion, the accelerator grid shows moderate corrosion, and the screen grid shows the least amount of corrosion. With the serious corrosion of the grids in 5 kW operation mode, the intercept current of the acceleration and deceleration grids increases substantially. Meanwhile, the cold gap between the accelerator grid and the screen grid decreases from 1 mm to 0.7 mm, while the cold gap between the accelerator grid and the decelerator grid increases from 1 mm to 1.25 mm after 1500 h of thruster operation. At equilibrium temperature with 5 kW power, the finite element method (FEM) simulation results show that the hot gap between the screen grid and the accelerator grid reduces to 0.2 mm. Accordingly, the hot gap between the accelerator grid and the decelerator grid increases to 1.5 mm. According to the fluid method, the plasma density simulated in most regions of the discharge chamber is 1 × 1018‑8 × 1018 m‑3. The upstream plasma density of the screen grid is in the range 6 × 1017‑6 × 1018 m‑3 and displays a parabolic characteristic. The charged particle tracing simulation method results show that the ion beam current without the thermal deformation of triple grids has optimal perveance status. The ion sputtering rates of the accelerator grid hole and the decelerator hole are 5.5 × 10‑14 kg s‑1 and

Effects of facility backpressure on the performance and plume of a Hall thruster

Science.gov (United States)

Walker, Mitchell Louis Ronald

2005-07-01

This dissertation presents research aimed at understanding the relationship between facility background pressure, Hall thruster performance, and plume characteristics. Due to the wide range of facilities used in Hall thruster testing, it is difficult for researchers to make adequate comparisons between data sets because of both dissimilar instrumentation and backpressures. The differences in the data sets are due to the ingestion of background gas into the Hall thruster discharge channel and charge-exchange collisions in the plume. Thus, this research aims to understand facility effects and to develop the tools needed to allow researchers to obtain relevant plume and performance data for a variety of chambers and backpressures. The first portion of this work develops a technique for calibrating a vacuum chamber in terms of pressure to account for elevated backpressures while testing Hall thrusters. Neutral gas background pressure maps of the Large Vacuum Test Facility are created at a series of cold anode flow rates and one hot flow rate at two UM/AFRL P5 5 kW Hall thruster operating conditions. These data show that a cold flow pressure map can be used to approximate the neutral background pressure in the chamber with the thruster in operation. In addition, the data are used to calibrate a numerical model that accurately predicts facility backpressure within a vacuum chamber of specified geometry and pumping speed. The second portion of this work investigates how facility backpressure influences the plume, plume diagnostics, and performance of the P5 Hall thruster. Measurements of the plume and performance characteristics over a wide range of pressures show that ingestion, a decrease in the downstream plasma potential, and broadening of the ion energy distribution function cause the increase in thrust with backpressure. Furthermore, a magnetically-filtered Faraday probe accurately measures ion current density at elevated operating pressures. The third portion of

Bi-Modal Micro-Cathode Arc Thruster for Cube Satellites

Science.gov (United States)

Chiu, Dereck

A new concept design, named the Bi-Modal Micro-Cathode Arc Thruster (BM-muCAT), has been introduced utilizing features from previous generations of muCATs and incorporating a multi-propellant functionality. This arc thruster is a micro-Newton level thruster based off of vacuum arc technology utilizing an enhanced magnetic field. Adjusting the magnetic field allows the thrusters performance to be varied. The goal of this thesis is to present a new generation of micro-cathode arc thrusters utilizing a bi-propellant, nickel and titanium, system. Three experimental procedures were run to test the new designs capabilities. Arc rotation experiment was used as a base experiment to ensure erosion was occurring uniformly along each electrode. Ion utilization efficiency was found, using an ion collector, to be up to 2% with the nickel material and 2.5% with the titanium material. Ion velocities were also studied using a time-of-flight method with an enhanced ion detection system. This system utilizes double electrostatic probes to measure plasma propagation. Ion velocities were measured to be 10km/s and 20km/s for nickel and titanium without a magnetic field. With an applied magnetic field of 0.2T, nickel ion velocities almost doubled to about 17km/s, while titanium ion velocities also increased to about 30km/s.

High-thrust and low-power operation of a 30-cm-diameter mercury ion thruster

Science.gov (United States)

Beattie, J. R.; Kami, S.

1981-01-01

An investigation of a 30-cm-diameter mercury ion thruster designed for high-thrust and low-power operation is described. Experimental results are presented which indicate that good performance and long lifetime are achieved by using a boundary magnetic field arrangement to confine the ionizing electrons. Details of advanced ion-optics designs are discussed, and performance measurements obtained with an advanced two-grid ion-optics assembly are presented. Scaling of the state-of-the-art hollow cathode for higher emission-current capability is described, and performance and lifetime measurements are presented for the scaled cathode.

Temperature Gradient in Hall Thrusters

International Nuclear Information System (INIS)

Staack, D.; Raitses, Y.; Fisch, N.J.

2003-01-01

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 significantly smaller than might be expected by models employing Ohmic heating of electrons

Near-Surface Plasma Characterization of the 12.5-kW NASA TDU1 Hall Thruster

Science.gov (United States)

Shastry, Rohit; Huang, Wensheng; Kamhawi, Hani

2015-01-01

To advance the state-of-the-art in Hall thruster technology, NASA is developing a 12.5-kW, high-specific-impulse, high-throughput thruster for the Solar Electric Propulsion Technology Demonstration Mission. In order to meet the demanding lifetime requirements of potential missions such as the Asteroid Redirect Robotic Mission, magnetic shielding was incorporated into the thruster design. Two units of the resulting thruster, called the Hall Effect Rocket with Magnetic Shielding (HERMeS), were fabricated and are presently being characterized. The first of these units, designated the Technology Development Unit 1 (TDU1), has undergone extensive performance and thermal characterization at NASA Glenn Research Center. A preliminary lifetime assessment was conducted by characterizing the degree of magnetic shielding within the thruster. This characterization was accomplished by placing eight flush-mounted Langmuir probes within each discharge channel wall and measuring the local plasma potential and electron temperature at various axial locations. Measured properties indicate a high degree of magnetic shielding across the throttle table, with plasma potential variations along each channel wall being less than or equal to 5 eV and electron temperatures being maintained at less than or equal to 5 eV, even at 800 V discharge voltage near the thruster exit plane. These properties indicate that ion impact energies within the HERMeS will not exceed 26 eV, which is below the expected sputtering threshold energy for boron nitride. Parametric studies that varied the facility backpressure and magnetic field strength at 300 V, 9.4 kW, illustrate that the plasma potential and electron temperature are insensitive to these parameters, with shielding being maintained at facility pressures 3X higher and magnetic field strengths 2.5X higher than nominal conditions. Overall, the preliminary lifetime assessment indicates a high degree of shielding within the HERMeS TDU1, effectively

High thrust-to-power ratio micro-cathode arc thruster

Directory of Open Access Journals (Sweden)

Joseph Lukas

2016-02-01

Full Text Available The Micro-Cathode Arc Thruster (μCAT is an electric propulsion device that ablates solid cathode material, through an electrical vacuum arc discharge, to create plasma and ultimately produce thrust in the μN to mN range. About 90% of the arc discharge current is conducted by electrons, which go toward heating the anode and contribute very little to thrust, with only the remaining 10% going toward thrust in the form of ion current. A preliminary set of experiments were conducted to show that, at the same power level, thrust may increase by utilizing an ablative anode. It was shown that ablative anode particles were found on a collection plate, compared to no particles from a non-ablative anode, while another experiment showed an increase in ion-to-arc current by approximately 40% at low frequencies compared to the non-ablative anode. Utilizing anode ablation leads to an increase in thrust-to-power ratio in the case of the μCAT.

Concept Study of Radio Frequency (RF Plasma Thruster for Space Propulsion

Directory of Open Access Journals (Sweden)

Anna-Maria Theodora ANDREESCU

2016-12-01

Full Text Available Electric thrusters are capable of accelerating ions to speeds that are impossible to reach using chemical reaction. Recent advances in plasma-based concepts have led to the identification of electromagnetic (RF generation and acceleration systems as able to provide not only continuous thrust, but also highly controllable and wide-range exhaust velocities. For Future Space Propulsion there is a pressing need for low pressure, high mass flow rate and controlled ion energies. This paper explores the potential of using RF heated plasmas for space propulsion in order to mitigate the electric propulsion problems caused by erosion and gain flexibility in plasma manipulation. The main key components of RF thruster architecture are: a feeding system able to provide the required neutral gas flow, plasma source chamber, antenna/electrodes wrapped around the discharge tube and optimized electromagnetic field coils for plasma confinement. A preliminary analysis of system performance (thrust, specific impulse, efficiency is performed along with future plans of Space Propulsion based on this new concept of plasma mechanism.

Microwave Discharge Ion Engines onboard Hayabusa Asteroid Explorer

International Nuclear Information System (INIS)

Kuninaka, Hitoshi

2008-01-01

The Hayabusa spacecraft rendezvoused with the asteroid Itokawa in 2005 after the powered flight in the deep space by the μl0 cathode-less electron cyclotron resonance ion engines. Though the spacecraft was seriously damaged after the successful soft-landing and lift-off, the xenon cold gas jets from the ion engines rescued it. New attitude stabilization method using a single reaction wheel, the ion beam jets, and the photon pressure was established and enabled the homeward journey from April 2007 aiming the Earth return on 2010. The total accumulated operational time of the ion engines reaches 31,400 hours at the end of 2007. One of four thrusters achieved 13,400-hour space operation

A structural and thermal packaging approach for power processing units for 30-cm ion thrusters

Science.gov (United States)

Maloy, J. E.; Sharp, G. R.

1975-01-01

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.

Diagnostics Systems for Permanent Hall Thrusters Development

Science.gov (United States)

Ferreira, Jose Leonardo; Soares Ferreira, Ivan; Santos, Jean; Miranda, Rodrigo; Possa, M. Gabriela

This work describes the development of Permanent Magnet Hall Effect Plasma Thruster (PHALL) and its diagnostic systems at The Plasma Physics Laboratory of University of Brasilia. The project consists on the construction and characterization of plasma propulsion engines based on the Hall Effect. Electric thrusters have been employed in over 220 successful space missions. Two types stand out: the Hall-Effect Thruster (HET) and the Gridded Ion Engine (GIE). The first, which we deal with in this project, has the advantage of greater simplicity of operation, a smaller weight for the propulsion subsystem and a longer shelf life. It can operate in two configurations: magnetic layer and anode layer, the difference between the two lying in the positioning of the anode inside the plasma channel. A Hall-Effect Thruster-HET is a type of plasma thruster in which the propellant gas is ionized and accelerated by a magneto hydrodynamic effect combined with electrostatic ion acceleration. So the essential operating principle of the HET is that it uses a J x B force and an electrostatic potential to accelerate ions up to high speeds. In a HET, the attractive negative charge is provided by electrons at the open end of the Thruster instead of a grid, as in the case of the electrostatic ion thrusters. A strong radial magnetic field is used to hold the electrons in place, with the combination of the magnetic field and the electrostatic potential force generating a fast circulating electron current, the Hall current, around the axis of the Thruster, mainly composed by drifting electrons in an ion plasma background. Only a slow axial drift towards the anode occurs. The main attractive features of the Hall-Effect Thruster are its simple design and operating principles. Most of the Hall-Effect Thrusters use electromagnet coils to produce the main magnetic field responsible for plasma generation and acceleration. In this paper we present a different new concept, a Permanent Magnet Hall

Operation of a Segmented Hall Thruster with Low-sputtering Carbon-velvet Electrodes

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Dunaevsky, A.; Fisch, N.J.

2005-01-01

Carbon fiber velvet material provides exceptional sputtering resistance properties exceeding those for graphite and carbon composite materials. A 2 kW Hall thruster with segmented electrodes made of this material was operated in the discharge voltage range of 200-700 V. The arcing between the floating velvet electrodes and the plasma was visually observed, especially, during the initial conditioning time, which lasted for about 1 h. The comparison of voltage versus current and plume characteristics of the Hall thruster with and without segmented electrodes indicates that the magnetic insulation of the segmented thruster improves with the discharge voltage at a fixed magnetic field. The observations reported here also extend the regimes wherein the segmented Hall thruster can have a narrower plume than that of the conventional nonsegmented thruster

Determination of the Hall Thruster Operating Regimes

International Nuclear Information System (INIS)

L. Dorf; V. Semenov; Y. Raitses; N.J. Fisch

2002-04-01

A quasi one-dimensional (1-D) steady-state model of the Hall thruster is presented. For the same discharge voltage two operating regimes are possible -- with and without the anode sheath. For given mass flow rate, magnetic field profile and discharge voltage a unique solution can be constructed, assuming that the thruster operates in one of the regimes. However, we show that for a given temperature profile the applied discharge voltage uniquely determines the operating regime: for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. It is also shown that a good correlation between the quasi 1-D model and experimental results can be achieved by selecting an appropriate electron mobility and temperature profile

Effect of Anode Dielectric Coating on Hall Thruster Operation

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.; Semenov, V.

2003-01-01

An interesting phenomenon observed in the near-anode region of a Hall thruster is that the anode fall changes from positive to negative upon removal of the dielectric coating, which is produced on the anode surface during the normal course of Hall thruster operation. The anode fall might affect the thruster lifetime and acceleration efficiency. The effect of the anode coating on the anode fall is studied experimentally using both biased and emissive probes. Measurements of discharge current oscillations indicate that thruster operation is more stable with the coated anode

Magnesium Hall Thruster

Science.gov (United States)

Szabo, James J.

2015-01-01

This Phase II project is developing a magnesium (Mg) Hall effect thruster system that would open the door for in situ resource utilization (ISRU)-based solar system exploration. Magnesium is light and easy to ionize. For a Mars- Earth transfer, the propellant mass savings with respect to a xenon Hall effect thruster (HET) system are enormous. Magnesium also can be combusted in a rocket with carbon dioxide (CO2) or water (H2O), enabling a multimode propulsion system with propellant sharing and ISRU. In the near term, CO2 and H2O would be collected in situ on Mars or the moon. In the far term, Mg itself would be collected from Martian and lunar regolith. In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was developed and tested. Controlled, steady operation at constant voltage and power was demonstrated. Preliminary measurements indicate a specific impulse (Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The feasibility of delivering fluidized Mg powder to a medium- or high-power thruster also was demonstrated. Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers improved the medium power thruster system and characterized it in detail. Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A fluidized powder feed system supporting the high-power thruster was built and delivered to Busek Company, Inc.

A data acquisition and storage system for the ion auxiliary propulsion system cyclic thruster test

Science.gov (United States)

Hamley, John A.

1989-01-01

A nine-track tape drive interfaced to a standard personal computer was used to transport data from a remote test site to the NASA Lewis mainframe computer for analysis. The Cyclic Ground Test of the Ion Auxiliary Propulsion System (IAPS), which successfully achieved its goal of 2557 cycles and 7057 hr of thrusting beam on time generated several megabytes of test data over many months of continuous testing. A flight-like controller and power supply were used to control the thruster and acquire data. Thruster data was converted to RS232 format and transmitted to a personal computer, which stored the raw digital data on the nine-track tape. The tape format was such that with minor modifications, mainframe flight data analysis software could be used to analyze the Cyclic Ground Test data. The personal computer also converted the digital data to engineering units and displayed real time thruster parameters. Hardcopy data was printed at a rate dependent on thruster operating conditions. The tape drive provided a convenient means to transport the data to the mainframe for analysis, and avoided a development effort for new data analysis software for the Cyclic test. This paper describes the data system, interfacing and software requirements.

In-Situ Measurement of Hall Thruster Erosion Using a Fiber Optic Regression Probe

Science.gov (United States)

Polzin, Kurt; Korman, Valentin

2009-01-01

One potential life-limiting mechanism in a Hall thruster is the erosion of the ceramic material comprising the discharge channel. This is especially true for missions that require long thrusting periods and can be problematic for lifetime qualification, especially when attempting to qualify a thruster by analysis rather than a test lasting the full duration of the mission. In addition to lifetime, several analytical and numerical models include electrode erosion as a mechanism contributing to enhanced transport properties. However, there is still a great deal of dispute over the importance of erosion to transport in Hall thrusters. The capability to perform an in-situ measurement of discharge channel erosion is useful in addressing both the lifetime and transport concerns. An in-situ measurement would allow for real-time data regarding the erosion rates at different operating points, providing a quick method for empirically anchoring any analysis geared towards lifetime qualification. Erosion rate data over a thruster s operating envelope would also be useful in the modeling of the detailed physics inside the discharge chamber. There are many different sensors and techniques that have been employed to quantify discharge channel erosion in Hall thrusters. Snapshots of the wear pattern can be obtained at regular shutdown intervals using laser profilometry. Many non-intrusive techniques of varying complexity and sensitivity have been employed to detect the time-varying presence of erosion products in the thruster plume. These include the use quartz crystal microbalances, emission spectroscopy, laser induced flourescence, and cavity ring-down spectroscopy. While these techniques can provide a very accurate picture of the level of eroded material in the thruster plume, it is more difficult to use them to determine the location from which the material was eroded. Furthermore, none of the methods cited provide a true in-situ measure of erosion at the channel surface while

Electromagnetic Spacecraft Propulsion Motor and a Permanent Magnet (PM-Drive) Thruster

Science.gov (United States)

Ahmadov, B. A.

2018-04-01

Ion thrusters are designed to be used for realization of a Mars Sample Return mission. The competing technologies with ion thrusters are electromagnetic spacecraft propulsion motors. I'm an engineer and engage in the creation of the new electromagnetic propulsion motors.

Determination of the Hall Thruster Operating Regimes; TOPICAL

International Nuclear Information System (INIS)

L. Dorf; V. Semenov; Y. Raitses; N.J. Fisch

2002-01-01

A quasi one-dimensional (1-D) steady-state model of the Hall thruster is presented. For the same discharge voltage two operating regimes are possible - with and without the anode sheath. For given mass flow rate, magnetic field profile and discharge voltage a unique solution can be constructed, assuming that the thruster operates in one of the regimes. However, we show that for a given temperature profile the applied discharge voltage uniquely determines the operating regime: for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. It is also shown that a good correlation between the quasi 1-D model and experimental results can be achieved by selecting an appropriate electron mobility and temperature profile

Design and Testing of a Hall Effect Thruster with Additively Manufactured Components

Science.gov (United States)

Hopping, Ethan

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville to study the application of low-cost additive manufacturing in the design and fabrication of Hall thrusters. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. The thruster features channel walls and a propellant distributor that were manufactured using 3D printing with a variety of materials including ABS, ULTEM, and glazed ceramic. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. The design of the thruster and the transient performance measurements are presented here. Measured thrust ranged from 17.2 mN to 30.4 mN over a discharge power of 280 W to 520 W with an anode Isp range of 870 s to 1450 s. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state. While the current thruster design is not yet ready for continuous operation, revisions to the device that could enable longer duration tests are discussed.

A Small Modular Laboratory Hall Effect Thruster

Science.gov (United States)

Lee, Ty Davis

Electric propulsion technologies promise to revolutionize access to space, opening the door for mission concepts unfeasible by traditional propulsion methods alone. The Hall effect thruster is a relatively high thrust, moderate specific impulse electric propulsion device that belongs to the class of electrostatic thrusters. Hall effect thrusters benefit from an extensive flight history, and offer significant performance and cost advantages when compared to other forms of electric propulsion. Ongoing research on these devices includes the investigation of mechanisms that tend to decrease overall thruster efficiency, as well as the development of new techniques to extend operational lifetimes. This thesis is primarily concerned with the design and construction of a Small Modular Laboratory Hall Effect Thruster (SMLHET), and its operation on argon propellant gas. Particular attention was addressed at low-cost, modular design principles, that would facilitate simple replacement and modification of key thruster parts such as the magnetic circuit and discharge channel. This capability is intended to facilitate future studies of device physics such as anomalous electron transport and magnetic shielding of the channel walls, that have an impact on thruster performance and life. Preliminary results demonstrate SMLHET running on argon in a manner characteristic of Hall effect thrusters, additionally a power balance method was utilized to estimate thruster performance. It is expected that future thruster studies utilizing heavier though more expensive gases like xenon or krypton, will observe increased efficiency and stability.

Plasma Characterization of Hall Thruster with Active and Passive Segmented Electrodes

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Fisch, N.J.

2002-01-01

Non-emissive electrodes and ceramic spacers placed along the Hall thruster channel are shown to affect the plasma potential distribution and the thruster operation. These effects are associated with physical properties of the electrode material and depend on the electrode configuration, geometry and the magnetic field distribution. An emissive segmented electrode was able to maintain thruster operation by supplying an additional electron flux to sustain the plasma discharge between the anode and cathode neutralizer. These results indicate the possibility of new configurations for segmented electrode Hall thruster

Hall Thruster Modeling with a Given Temperature Profile

International Nuclear Information System (INIS)

Dorf, L.; Semenov, V.; Raitses, Y.; Fisch, N.J.

2002-01-01

A quasi one-dimensional steady-state model of the Hall thruster is presented. For given mass flow rate, magnetic field profile, and discharge voltage the unique solution can be constructed, assuming that the thruster operates in one of the two regimes: with or without the anode sheath. It is shown that for a given temperature profile, the applied discharge voltage uniquely determines the operating regime; for discharge voltages greater than a certain value, the sheath disappears. That result is obtained over a wide range of incoming neutral velocities, channel lengths and widths, and cathode plane locations. A good correlation between the quasi one-dimensional model and experimental results can be achieved by selecting an appropriate temperature profile. We also show how the presented model can be used to obtain a two-dimensional potential distribution

Retrofit and acceptance test of 30-cm ion thrusters

Science.gov (United States)

Poeschel, R. L.

1981-01-01

Six 30 cm mercury thrusters were modified to the J-series design and evaluated using standardized test procedures. The thruster performance meets the design objectives (lifetime objective requires verification), and documentation (drawings, etc.) for the design is completed and upgraded. The retrofit modifications are described and the test data for the modifications are presented and discussed.

Magnetically enhanced vacuum arc thruster

International Nuclear Information System (INIS)

Keidar, Michael; Schein, Jochen; Wilson, Kristi; Gerhan, Andrew; Au, Michael; Tang, Benjamin; Idzkowski, Luke; Krishnan, Mahadevan; Beilis, Isak I

2005-01-01

A hydrodynamic model of the vacuum arc thruster and its plume is described. Primarily an effect of the magnetic field on the plume expansion and plasma generation is considered. Two particular examples are investigated, namely the magnetically enhanced co-axial vacuum arc thruster (MVAT) and the vacuum arc thruster with ring electrodes (RVAT). It is found that the magnetic field significantly decreases the plasma plume radial expansion under typical conditions. Predicted plasma density profiles in the plume of the MVAT are compared with experimental profiles, and generally a good agreement is found. In the case of the RVAT the influence of the magnetic field leads to plasma jet deceleration, which explains the non-monotonic dependence of the ion current density, on an axial magnetic field observed experimentally

Magnetically enhanced vacuum arc thruster

Energy Technology Data Exchange (ETDEWEB)

Keidar, Michael [University of Michigan, Ann Arbor 48109 MI (United States); Schein, Jochen [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Wilson, Kristi [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Gerhan, Andrew [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Au, Michael [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Tang, Benjamin [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Idzkowski, Luke [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Krishnan, Mahadevan [Alameda Applied Science Corporation, San Leandro, CA 94577 (United States); Beilis, Isak I [Tel Aviv University, Tel Aviv (Israel)

2005-11-01

A hydrodynamic model of the vacuum arc thruster and its plume is described. Primarily an effect of the magnetic field on the plume expansion and plasma generation is considered. Two particular examples are investigated, namely the magnetically enhanced co-axial vacuum arc thruster (MVAT) and the vacuum arc thruster with ring electrodes (RVAT). It is found that the magnetic field significantly decreases the plasma plume radial expansion under typical conditions. Predicted plasma density profiles in the plume of the MVAT are compared with experimental profiles, and generally a good agreement is found. In the case of the RVAT the influence of the magnetic field leads to plasma jet deceleration, which explains the non-monotonic dependence of the ion current density, on an axial magnetic field observed experimentally.

Thermal-environmental testing of a 30-cm engineering model thruster

Science.gov (United States)

Mirtich, M. J.

1976-01-01

An experimental test program was carried out to document all 30-cm electron bombardment Hg ion bombardment thruster functions and characteristics over the thermal environment of several proposed missions. An engineering model thruster was placed in a thermal test facility equipped with -196 C walls and solar simulation. The thruster was cold soaked and exposed to simulated eclipses lasting in duration from 17 to 72 minutes. The thruster was operated at quarter, to full beam power in various thermal configurations which simulated multiple thruster operation, and was also exposed to 1 and 2 suns solar simulation. Thruster control characteristics and constraints; performance, including thrust magnitude and direction; and structural integrity were evaluated over the range of thermal environments tested.

Hall Thruster Thermal Modeling and Test Data Correlation

Science.gov (United States)

Myers, James; Kamhawi, Hani; Yim, John; Clayman, Lauren

2016-01-01

The life of Hall Effect thrusters are primarily limited by plasma erosion and thermal related failures. NASA Glenn Research Center (GRC) in cooperation with the Jet Propulsion Laboratory (JPL) have recently completed development of a Hall thruster with specific emphasis to mitigate these limitations. Extending the operational life of Hall thursters makes them more suitable for some of NASA's longer duration interplanetary missions. This paper documents the thermal model development, refinement and correlation of results with thruster test data. Correlation was achieved by minimizing uncertainties in model input and recognizing the relevant parameters for effective model tuning. Throughout the thruster design phase the model was used to evaluate design options and systematically reduce component temperatures. Hall thrusters are inherently complex assemblies of high temperature components relying on internal conduction and external radiation for heat dispersion and rejection. System solutions are necessary in most cases to fully assess the benefits and/or consequences of any potential design change. Thermal model correlation is critical since thruster operational parameters can push some components/materials beyond their temperature limits. This thruster incorporates a state-of-the-art magnetic shielding system to reduce plasma erosion and to a lesser extend power/heat deposition. Additionally a comprehensive thermal design strategy was employed to reduce temperatures of critical thruster components (primarily the magnet coils and the discharge channel). Long term wear testing is currently underway to assess the effectiveness of these systems and consequently thruster longevity.

Los Alamos NEP research in advanced plasma thrusters

Science.gov (United States)

Schoenberg, Kurt; Gerwin, Richard

1991-01-01

Research was initiated in advanced plasma thrusters that capitalizes on lab capabilities in plasma science and technology. The goal of the program was to examine the scaling issues of magnetoplasmadynamic (MPD) thruster performance in support of NASA's MPD thruster development program. The objective was to address multi-megawatt, large scale, quasi-steady state MPD thruster performance. Results to date include a new quasi-steady state operating regime which was obtained at space exploration initiative relevant power levels, that enables direct coaxial gun-MPD comparisons of thruster physics and performance. The radiative losses are neglible. Operation with an applied axial magnetic field shows the same operational stability and exhaust plume uniformity benefits seen in MPD thrusters. Observed gun impedance is in close agreement with the magnetic Bernoulli model predictions. Spatial and temporal measurements of magnetic field, electric field, plasma density, electron temperature, and ion/neutral energy distribution are underway. Model applications to advanced mission logistics are also underway.

A high sensitivity momentum flux measuring instrument for plasma thruster exhausts and diffusive plasmas.

Science.gov (United States)

West, Michael D; Charles, Christine; Boswell, Rod W

2009-05-01

A high sensitivity momentum flux measuring instrument based on a compound pendulum has been developed for use with electric propulsion devices and radio frequency driven plasmas. A laser displacement system, which builds upon techniques used by the materials science community for surface stress measurements, is used to measure with high sensitivity the displacement of a target plate placed in a plasma thruster exhaust. The instrument has been installed inside a vacuum chamber and calibrated via two different methods and is able to measure forces in the range of 0.02-0.5 mN with a resolution of 15 microN. Measurements have been made of the force produced from the cold gas flow and with a discharge ignited using argon propellant. The plasma is generated using a Helicon Double Layer Thruster prototype. The instrument target is placed about 1 mean free path for ion-neutral charge exchange collisions downstream of the thruster exit. At this position, the plasma consists of a low density ion beam (10%) and a much larger downstream component (90%). The results are in good agreement with those determined from the plasma parameters measured with diagnostic probes. Measurements at various flow rates show that variations in ion beam velocity and plasma density and the resulting momentum flux can be measured with this instrument. The instrument target is a simple, low cost device, and since the laser displacement system used is located outside the vacuum chamber, the measurement technique is free from radio frequency interference and thermal effects. It could be used to measure the thrust in the exhaust of other electric propulsion devices and the momentum flux of ion beams formed by expanding plasmas or fusion experiments.

Integrated Stirling Convertor and Hall Thruster Test Conducted

Science.gov (United States)

Mason, Lee S.

2002-01-01

An important aspect of implementing Stirling Radioisotope Generators on future NASA missions is the integration of the generator and controller with potential spacecraft loads. Some recent studies have indicated that the combination of Stirling Radioisotope Generators and electric propulsion devices offer significant trip time and payload fraction benefits for deep space missions. A test was devised to begin to understand the interactions between Stirling generators and electric thrusters. An electrically heated RG- 350 (350-W output) Stirling convertor, designed and built by Stirling Technology Company of Kennewick, Washington, under a NASA Small Business Innovation Research agreement, was coupled to a 300-W SPT-50 Hall-effect thruster built for NASA by the Moscow Aviation Institute (RIAME). The RG-350 and the SPT-50 shown, were installed in adjacent vacuum chamber ports at NASA Glenn Research Center's Electric Propulsion Laboratory, Vacuum Facility 8. The Stirling electrical controller interfaced directly with the Hall thruster power-processing unit, both of which were located outside of the vacuum chamber. The power-processing unit accepted the 48 Vdc output from the Stirling controller and distributed the power to all the loads of the SPT-50, including the magnets, keeper, heater, and discharge. On February 28, 2001, the Glenn test team successfully operated the Hall-effect thruster with the Stirling convertor. This is the world's first known test of a dynamic power source with electric propulsion. The RG-350 successfully managed the transition from the purely resistive load bank within the Stirling controller to the highly capacitive power-processing unit load. At the time of the demonstration, the Stirling convertor was operating at a hot temperature of 530 C and a cold temperature of -6 C. The linear alternator was producing approximately 250 W at 109 Vac, while the power-processing unit was drawing 175 W at 48 Vdc. The majority of power was delivered to the

The influence of magnetic field strength in ionization stage on ion transport between two stages of a double stage Hall thruster

International Nuclear Information System (INIS)

Yu Daren; Song Maojiang; Li Hong; Liu Hui; Han Ke

2012-01-01

It is futile for a double stage Hall thruster to design a special ionization stage if the ionized ions cannot enter the acceleration stage. Based on this viewpoint, the ion transport under different magnetic field strengths in the ionization stage is investigated, and the physical mechanisms affecting the ion transport are analyzed in this paper. With a combined experimental and particle-in-cell simulation study, it is found that the ion transport between two stages is chiefly affected by the potential well, the potential barrier, and the potential drop at the bottom of potential well. With the increase of magnetic field strength in the ionization stage, there is larger plasma density caused by larger potential well. Furthermore, the potential barrier near the intermediate electrode declines first and then rises up while the potential drop at the bottom of potential well rises up first and then declines as the magnetic field strength increases in the ionization stage. Consequently, both the ion current entering the acceleration stage and the total ion current ejected from the thruster rise up first and then decline as the magnetic field strength increases in the ionization stage. Therefore, there is an optimal magnetic field strength in the ionization stage to guide the ion transport between two stages.

Ion source based on Penning discharge for production of doubly charged helium ions

Directory of Open Access Journals (Sweden)

V. I. Voznyi

2017-11-01

Full Text Available The article presents the results of operation of ion source with Penning discharge developed in the IAP of NAS of Ukraine to produce doubly charged helium ions He2+ beam and to increase the energy of accelerated ions up to 3.2 MeV. This energy is necessary for ERDA channel when measuring hydrogen concentration in the structural materials used in nuclear engineering. The ion source parameters are the following: discharge voltage is 6 kV, discharge current is 0.8 - 1.2 mA, the current of singly charged helium ions He+ 24 μA, the current of doubly charged helium ions He2+ 0.5 μA.

Negative Ions in low pressure discharges

NARCIS (Netherlands)

Stoffels - Adamowicz, E.; Stoffels, W.W.; Vender, D.; Haverlag, M.; Kroesen, G.M.W.; Hoog, de F.J.

1995-01-01

Several aspects of negative ions in low pressure discharges are treated. The elementary processes, in which negative ions are produced and destroyed, are summarized. The influence of negative ions on plasma operation is analyzed in terms of transport equations. It is shown that diffusion, electric

ExB Measurements of a 200 W Xenon Hall Thruster (Preprint)

National Research Council Canada - National Science Library

Ekholm, Jared M; Hargus, Jr, William A

2007-01-01

Angularly resolved ion species fractions of Xe+1, Xe+2, and Xe+3 in a low power xenon Hall thruster Busek BHT-200 plume were measured using an ExB probe under a variety of thruster operating conditions and background pressures...

Electron Cross-field Transport in a Miniaturized Cylindrical Hall Thruster

International Nuclear Information System (INIS)

Smirnov Artem; Raitses Yevgeny; Fisch Nathaniel J

2005-01-01

Conventional annular Hall thrusters become inefficient when scaled to low power. Cylindrical Hall thrusters, which have lower surface-to-volume ratio, are more promising for scaling down. They presently exhibit performance comparable with conventional annular Hall thrusters. The present paper gives a review of the experimental and numerical investigations of electron crossfield transport in the 2.6 cm miniaturized cylindrical Hall thruster (100 W power level). We show that, in order to explain the discharge current observed for the typical operating conditions, the electron anomalous collision frequency ν b has to be on the order of the Bohm value, ν B ∼ ω c /16. The contribution of electron-wall collisions to cross-field transport is found to be insignificant. The optimal regimes of thruster operation at low background pressure (below 10 -5 Torr) in the vacuum tank appear to be different from those at higher pressure (∼ 10 -4 Torr)

Ions mobilities in corona discharge

International Nuclear Information System (INIS)

Bakhtaev, Sh. A.; Bochkareva, G. V.; Sydykova, G. K.

2000-01-01

Ion mobility in unipolar corona at small inter-electron distances (up to 0.01 m) when as coroning element serves micro-wire is consider. Experimental data of ion mobility in corona discharge external zone in atmospheric air are obtained and its comparative analysis with known data is worked out. (author)

Design and Testing of a Hall Effect Thruster with 3D Printed Channel and Propellant Distributor

Science.gov (United States)

Hopping, Ethan P.; Xu, Kunning G.

2017-01-01

The UAH-78AM is a low-power Hall effect thruster developed at the University of Alabama in Huntsville with channel walls and a propellant distributor manufactured using 3D printing. The goal of this project is to assess the feasibility of using unconventional materials to produce a low-cost functioning Hall effect thruster and consider how additive manufacturing can expand the design space and provide other benefits. A version of the thruster was tested at NASA Glenn Research Center to obtain performance metrics and to validate the ability of the thruster to produce thrust and sustain a discharge. An overview of the thruster design and transient performance measurements are presented here. Measured thrust ranged from 17.2 millinewtons to 30.4 millinewtons over a discharge power of 280 watts to 520 watts with an anode I (sub SP)(Specific Impulse) range of 870 seconds to 1450 seconds. Temperature limitations of materials used for the channel walls and propellant distributor limit the ability to run the thruster at thermal steady-state.

Maximizing Ion Current by Space Charge Neutralization using Negative Ions and Dust Particles

International Nuclear Information System (INIS)

Smirnov, A.; Raitses, Y.; Fisch, N.J.

2005-01-01

Ion current extracted from an ion source (ion thruster) can be increased above the Child-Langmuir limit if the ion space charge is neutralized. Similarly, the limiting kinetic energy density of the plasma flow in a Hall thruster might be exceeded if additional mechanisms of space charge neutralization are introduced. Space charge neutralization with high-mass negative ions or negatively charged dust particles seems, in principle, promising for the development of a high current or high energy density source of positive light ions. Several space charge neutralization schemes that employ heavy negatively charged particles are considered. It is shown that the proposed neutralization schemes can lead, at best, only to a moderate but nonetheless possibly important increase of the ion current in the ion thruster and the thrust density in the Hall thruster

A mechanical, thermal and electrical packaging design for a prototype power management and control system for the 30 cm mercury ion thruster

Science.gov (United States)

Sharp, G. R.; Gedeon, L.; Oglebay, J. C.; Shaker, F. S.; Siegert, C. E.

1978-01-01

A prototype electric power management and thruster control system for a 30 cm ion thruster is described. The system meets all of the requirements necessary to operate a thruster in a fully automatic mode. Power input to the system can vary over a full two to one dynamic range (200 to 400 V) for the solar array or other power source. The power management and control system is designed to protect the thruster, the flight system and itself from arcs and is fully compatible with standard spacecraft electronics. The system is easily integrated into flight systems which can operate over a thermal environment ranging from 0.3 to 5 AU. The complete power management and control system measures 45.7 cm (18 in.) x 15.2 cm (6 in.) x 114.8 cm (45.2 in.) and weighs 36.2 kg (79.7 lb). At full power the overall efficiency of the system is estimated to be 87.4 percent. Three systems are currently being built and a full schedule of environmental and electrical testing is planned.

A mechanical, thermal and electrical packaging design for a prototype power management and control system for the 30 cm mercury ion thruster

Science.gov (United States)

Sharp, G. R.; Gedeon, L.; Oglebay, J. C.; Shaker, F. S.; Siegert, C. E.

1978-01-01

A prototype Electric Power Management and Thruster Control System for a 30 cm ion thruster has been built and is ready to support a first mission application. The system meets all of the requirements necessary to operate a thruster in a fully automatic mode. Power input to the system can vary over a full two to one dynamic range (200 to 400 V) for the solar array or other power source. The Power Management and Control system is designed to protect the thruster, the flight system and itself from arcs and is fully compatible with standard spacecraft electronics. The system is designed to be easily integrated into flight systems which can operate over a thermal environment ranging from 0.3 to 5 AU. The complete Power Management and Control system measures 45.7 cm x 15.2 cm x 114.8 cm and weighs 36.2 kg. At full power the overall efficiency of the system is estimated to be 87.4 percent. Three systems are currently being built and a full schedule of environmental and electrical testing is planned.

Influence of discharge gap on the discharge stability in a short vacuum arc ion source

Energy Technology Data Exchange (ETDEWEB)

Chen, L. [Institute of Electronic Engineering, China Academy of Engineering Physics, Mianyang, Sichuan 621900 (China); Department of Engineering Physics, Tsinghua University, Beijing 100084 (China); Zhang, G. L.; Jin, D. Z.; Dai, J. Y. [Institute of Electronic Engineering, China Academy of Engineering Physics, Mianyang, Sichuan 621900 (China); Yang, L. [Institute of Nuclear Science and Technology, Louzhou University, Lanzhou, Gansu 730000 (China)

2012-02-15

The influence of the discharge gap between cathode and anode on the discharge stability in a short vacuum arc (SVA) ion source is presented in this paper. Planar cathode and cylindrical hollow anode made of titanium are investigated. There is a great need in present accelerator injection research for SVA source to produce the small deviation of the ion current beam. Current research shows that increasing the short discharge gap can reduce the level of ion current deviation and ion charge deviation from 29% and 31% to 15% and 17%, respectively. A microplasma plume generation mechanism in SVA and scanning electron microscopic results can be used to explain this interesting phenomenon.

Improvement of Flow Characteristics for an Advanced Plasma Thruster

International Nuclear Information System (INIS)

Inutake, M.; Hosokawa, Y.; Sato, R.; Ando, A.; Tobari, H.; Hattori, K.

2005-01-01

A higher specific impulse and a larger thrust are required for a manned interplanetary space thruster. Until the realization of a fusion-plasma thruster, a magneto-plasma-dynamic arcjet (MPDA) powered by a fission reactor is one of the promising candidates for a manned Mars space thruster. The MPDA plasma is accelerated axially by a self-induced j x B force. Thrust performance of the MPDA is expected to increase by applying a magnetic nozzle instead of a solid nozzle. In order to get a much higher thruster performance, two methods have been investigated in the HITOP device, Tohoku University. One is to use a magnetic Laval nozzle in the vicinity of the MPDA muzzle for converting the high ion thermal energy to the axial flow energy. The other is to heat ions by use of an ICRF antenna in the divergent magnetic nozzle. It is found that by use of a small-sized Laval-type magnetic nozzle, the subsonic flow near the muzzle is converted to be supersonic through the magnetic Laval nozzle. A fast-flowing plasma is successfully heated by use of an ICRF antenna in the magnetic beach configuration

Thermal modeling of cylindrical lithium ion battery during discharge cycle

International Nuclear Information System (INIS)

Jeon, Dong Hyup; Baek, Seung Man

2011-01-01

Highlights: → Transient and thermo-electric finite element analysis (FEA) of cylindrical lithium ion (Li-ion) battery was presented. → This model provides the thermal behavior of Li-ion battery during discharge cycle. → A LiCoO 2 /C battery at various discharge rates was investigated. → The contribution of heat source due to joule heating was significant at a high discharge rate. → The contribution of heat source due to entropy change was dominant at a low discharge rate. - Abstract: Transient and thermo-electric finite element analysis (FEA) of cylindrical lithium ion (Li-ion) battery was presented. The simplified model by adopting a cylindrical coordinate was employed. This model provides the thermal behavior of Li-ion battery during discharge cycle. The mathematical model solves conservation of energy considering heat generations due to both joule heating and entropy change. A LiCoO 2 /C battery at various discharge rates was investigated. The temperature profile from simulation had similar tendency with experiment. The temperature profile was decomposed with contributions of each heat sources and was presented at several discharge rates. It was found that the contribution of heat source due to joule heating was significant at a high discharge rate, whereas that due to entropy change was dominant at a low discharge rate. Also the effect of cooling condition and the LiNiCoMnO 2 /C battery were analyzed for the purpose of temperature reduction.

Modeling an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT)

Science.gov (United States)

Choi, Maria

2016-01-01

An iodine-operated 200-W Hall thruster plume has been simulated using a hybrid-PIC model to predict the spacecraft surface-plume interaction for spacecraft integration purposes. For validation of the model, the plasma potential, electron temperature, ion current flux, and ion number density of xenon propellant were compared with available measurement data at the nominal operating condition. To simulate iodine plasma, various collision cross sections were found and used in the model. While time-varying atomic iodine species (i.e., I, I+, I2+) information is provided by HPHall simulation at the discharge channel exit, the molecular iodine species (i.e., I2, I2+) are introduced as Maxwellian particles at the channel exit. Simulation results show that xenon and iodine plasma plumes appear to be very similar under the assumptions of the model. Assuming a sticking coefficient of unity, iodine deposition rate is estimated.

Performance of a Permanent-Magnet Cylindrical Hall-Effect Thruster

Science.gov (United States)

Polzin, K. A.; Sooby, E. S.; Kimberlin, A. C.; Raites, Y.; Merino, E.; Fisch, N. J.

2009-01-01

The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic topologies. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying higher thrust efficiency. Thruster performance measurements on this configuration were obtained over a power range of 70-350 W and with the cathode orifice located at three different axial positions relative to the thruster exit plane. The thrust levels over this power range were 1.25-6.5 mN, with anode efficiencies and specific impulses spanning 4-21% and 400-1950 s, respectively. The anode efficiency of the permanent-magnet thruster compares favorable with the efficiency of the electromagnet thruster when the power consumed by the electromagnets is taken into account.

Performance of a Cylindrical Hall-Effect Thruster Using Permanent Magnets

Science.gov (United States)

Polzin, Kurt A.; Raitses, Y.; Merino, E.; Fisch, N. J.

2009-01-01

While annular Hall thrusters can operate at high efficiency at kW power levels, it is difficult to construct one that operates over a broad envelope from 1 kW down to 100 W while maintaining an efficiency of 45-55%. Scaling to low power while holding the main dimensionless parameters constant requires a decrease in the thruster channel size and an increase in the magnetic field strength. Increasing the magnetic field becomes technically challenging since the field can saturate the miniaturized inner components of the magnetic circuit and scaling down the magnetic circuit leaves very little room for magnetic pole pieces and heat shields. In addition, the central magnetic pole piece defining the interior wall of the annular channel can experience excessive heat loads in a miniaturized Hall thruster, with the temperature eventually exceeding the Curie temperature of the material and in extreme circumstances leading to accelerated erosion of the channel wall. An alternative approach is to employ a cylindrical Hall thruster (CHT) geometry. Laboratory model CHTs have operated at power levels ranging from 50 W up to 1 kW. These thrusters exhibit performance characteristics that are comparable to conventional, annular Hall thrusters of similar size. Compared to the annular Hall thruster, the CHTs insulator surface area to discharge chamber volume ratio is lower. Consequently, there is the potential for reduced wall losses in the channel of a CHT, and any reduction in wall losses should translate into lower channel heating rates and reduced erosion, making the CHT geometry promising for low-power applications. This potential for high performance in the low-power regime has served as the impetus for research and development efforts aimed at understanding and improving CHT performance. Recently, a 2.6 cm channel diameter permanent magnet CHT (shown in Fig. 1) was tested. This thruster has the promise of reduced power consumption over previous CHT iterations that employed

Iodine Hall Thruster Propellant Feed System for a CubeSat

Science.gov (United States)

Polzin, Kurt A.

2014-01-01

There has been significant work recently in the development of iodine-fed Hall thrusters for in-space propulsion applications.1 The use of iodine as a propellant provides many advantages over present xenon-gas-fed Hall thruster systems. Iodine is a solid at ambient temperature (no pressurization required) and has no special handling requirements, making it safe for secondary flight opportunities. It has exceptionally high ?I sp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing system level advantages over mid-term high power electric propulsion options. Iodine provides thrust and efficiency that are comparable to xenonfed Hall thrusters while operating in the same discharge current and voltage regime, making it possible to leverage the development of flight-qualified xenon Hall thruster power processing units for the iodine application. Work at MSFC is presently aimed at designing, integrating, and demonstrating a flight-like iodine feed system suitable for the Hall thruster application. This effort represents a significant advancement in state-of-the-art. Though Iodine thrusters have demonstrated high performance with mission enabling potential, a flight-like feed system has never been demonstrated and iodine compatible components do not yet exist. Presented in this paper is the end-to-end integrated feed system demonstration. The system includes a propellant tank with active feedback-control heating, fill and drain interfaces, latching and proportional flow control valves (PFCV), flow resistors, and flight-like CubeSat power and control electronics. Hardware is integrated into a CubeSat-sized structure, calibrated and tested under vacuum conditions, and operated under under hot-fire conditions using a Busek BHT-200 thruster designed for iodine. Performance of the system is evaluated thorugh accurate measurement of thrust and a calibrated of mass flow rate measurement, which is a function of

Corona discharge ion mobility spectrometry at reduced pressures

International Nuclear Information System (INIS)

Tabrizchi, Mahmoud; Rouholahnejad, Fereshteh

2004-01-01

Ion mobility spectrometers (IMSs) normally operate at ambient pressure. In this work an IMS cell has been designed and constructed to allow the pressure to be reduced inside the IMS cell. In this cell, corona discharge was employed as the ionization source. Reducing pressure affected both the discharge and the performance of the IMS. The discharge current was observed to increase with reducing pressure while the ignition potential decreased. The ion current received at the collector plate was also increased about 50 times when the pressure was reduced from ambient pressure to 15 Torr. The higher ion current can lead to an extended dynamic range. IMS spectra were recorded at various pressures and the results show that the drift times shift perfectly linear with pressure. This suggests that unlike temperature, pressure correction for ion mobility spectra is as simple as multiplying the drift times by a factor of 760/P

Integration Tests of the 4 kW-class High Voltage Hall Accelerator Power Processing Unit with the HiVHAc and the SPT-140 Hall Effect Thrusters

Science.gov (United States)

Kamhawi, Hani; Pinero, Luis; Haag, Thomas; Huang, Wensheng; Ahern, Drew; Liang, Ray; Shilo, Vlad

2016-01-01

NASAs Science Mission Directorate is sponsoring the development of a 4 kW-class Hall propulsion system for implementation in NASA science and exploration missions. The main components of the system include the High Voltage Hall Accelerator (HiVHAc), an engineering model power processing unit (PPU) developed by Colorado Power Electronics, and a xenon flow control module (XFCM) developed by VACCO Industries. NASA Glenn Research Center is performing integrated tests of the Hall thruster propulsion system. This presentation presents results from integrated tests of the PPU and XFCM with the HiVHAc engineering development thruster and a SPT-140 thruster provided by Space System Loral. The results presented in this paper demonstrate thruster discharge initiation, open-loop and closed-loop control of the discharge current with anode flow for both the HiVHAc and the SPT-140 thrusters. Integrated tests with the SPT-140 thruster indicated that the PPU was able to repeatedly initiate the thrusters discharge, achieve steady state operation, and successfully throttle the thruster between 1.5 and 4.5 kW. The measured SPT-140 performance was identical to levels reported by Space Systems Loral.

Pulsed inductive thruster performance data base for megawatt-class engine applications

International Nuclear Information System (INIS)

Dailey, C.L.; Lovberg, R.H.

1993-01-01

The pulsed inductive thruster (PIT) is an electrodeless plasma accelerator employing a large (1m diameter) spiral coil energized by a capacitor bank discharge. The bank can be repetitively recharged by a nuclear electric generator for continuous MW level operation. The coil can be designed as a transformer that permits thruster operation at the generator voltage, which results in a low thruster specific mass. Specific impulse (I sp ) can be readily altered by changing the propellant valve plenum pressure. Performance curves generated from mesausred impulse, injected mass and capacitor bank energy are presented for argon, ammonia, hydrazine, carbon dioxide and helium. The highest performance measured to date is 48% efficiency at 4000 seconds I sp with ammonia. The development of a theoretical model of the thruster, which assumes a fully ionized plasma, is presented in an appendix

Control Valve for Miniature Xenon Ion Thruster, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — NASA is continuing its development of electric propulsion engines for various applications. Efforts have been directed toward both large and small thrusters,...

An evaluation of krypton propellant in Hall thrusters

Science.gov (United States)

Linnell, Jesse Allen

Due to its high specific impulse and low price, krypton has long sparked interest as an alternate Hall thruster propellant. Unfortunately at the moment, krypton's relatively poor performance precludes it as a legitimate option. This thesis presents a detailed investigation into krypton operation in Hall thrusters. These findings suggest that the performance gap can be decreased to 4% and krypton can finally become a realistic propellant option. Although krypton has demonstrated superior specific impulse, the xenon-krypton absolute efficiency gap ranges between 2 and 15%. A phenomenological performance model indicates that the main contributors to the efficiency gap are propellant utilization and beam divergence. Propellant utilization and beam divergence have relative efficiency deficits of 5 and 8%, respectively. A detailed characterization of internal phenomena is conducted to better understand the xenon-krypton efficiency gap. Krypton's large beam divergence is found to be related to a defocusing equipotential structure and a weaker magnetic field topology. Ionization processes are shown to be linked to the Hall current, the magnetic mirror topology, and the perpendicular gradient of the magnetic field. Several thruster design and operational suggestions are made to optimize krypton efficiency. Krypton performance is optimized for discharge voltages above 500 V and flow rates corresponding to an a greater than 0.015 mg/(mm-s), where alpha is a function of flow rate and discharge channel dimensions (alpha = m˙alphab/Ach). Performance can be further improved by increasing channel length or decreasing channel width for a given flow rate. Also, several magnetic field design suggestions are made to enhance ionization and beam focusing. Several findings are presented that improve the understanding of general Hall thruster physics. Excellent agreement is shown between equipotential lines and magnetic field lines. The trim coil is shown to enhance beam focusing

Mechanical Design of Carbon Ion Optics

Science.gov (United States)

Haag, Thomas

2005-01-01

Carbon Ion Optics are expected to provide much longer thruster life due to their resistance to sputter erosion. There are a number of different forms of carbon that have been used for fabricating ion thruster optics. The mechanical behavior of carbon is much different than that of most metals, and poses unique design challenges. In order to minimize mission risk, the behavior of carbon must be well understood, and components designed within material limitations. Thermal expansion of the thruster structure must be compatible with thermal expansion of the carbon ion optics. Specially designed interfaces may be needed so that grid gap and aperture alignment are not adversely affected by dissimilar material properties within the thruster. The assembled thruster must be robust and tolerant of launch vibration. The following paper lists some of the characteristics of various carbon materials. Several past ion optics designs are discussed, identifying strengths and weaknesses. Electrostatics and material science are not emphasized so much as the mechanical behavior and integration of grid electrodes into an ion thruster.

Low Frequency Plasma Oscillations in a 6-kW Magnetically Shielded Hall Thruster

Science.gov (United States)

Jorns, Benjamin A.; Hofery, Richard R.

2013-01-01

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.

Real-Tme Boron Nitride Erosion Measurements of the HiVHAc Thruster via Cavity Ring-Down Spectroscopy

Science.gov (United States)

Lee, Brian C.; Yalin, Azer P.; Gallimore, Alec; Huang, Wensheng; Kamhawi, Hani

2013-01-01

Cavity ring-down spectroscopy was used to make real-time erosion measurements from the NASA High Voltage Hall Accelerator thruster. The optical sensor uses 250 nm light to measure absorption of atomic boron in the plume of an operating Hall thruster. Theerosion rate of the High Voltage Hall Accelerator thruster was measured for discharge voltages ranging from 330 to 600 V and discharge powers ranging from 1 to 3 kW. Boron densities as high as 6.5 x 10(exp 15) per cubic meter were found within the channel. Using a very simple boronvelocity model, approximate volumetric erosion rates between 5.0 x 10(exp -12) and 8.2 x 10(exp -12) cubic meter per second were found.

Effects of non-Maxwellian electron velocity distribution function on two-stream instability in low-pressure discharges

International Nuclear Information System (INIS)

Sydorenko, D.; Smolyakov, A.; Kaganovich, I.; Raitses, Y.

2007-01-01

Electron emission from discharge chamber walls is important for plasma maintenance in many low-pressure discharges. The electrons emitted from the walls are accelerated by the sheath electric field and are injected into the plasma as an electron beam. Penetration of this beam through the plasma is subject to the two-stream instability, which tends to slow down the beam electrons and heat the plasma electrons. In the present paper, a one-dimensional particle-in-cell code is used to simulate these effects both in a collisionless plasma slab with immobile ions and in a cross-field discharge of a Hall thruster. The two-stream instability occurs if the total electron velocity distribution function of the plasma-beam system is a nonmonotonic function of electron speed. Low-pressure plasmas can be depleted of electrons with energy above the plasma potential. This study reveals that under such conditions the two-stream instability depends crucially on the velocity distribution function of electron emission. It is shown that propagation of the secondary electron beams in Hall thrusters may be free of the two-stream instability if the velocity distribution of secondary electron emission is a monotonically decaying function of speed. In this case, the beams propagate between the walls with minimal loss of the beam current and the secondary electron emission does not affect the thruster plasma properties

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Science.gov (United States)

Martin, Adam K.; Polzin, Kurt A.; Eskridge, Richard H.; Smith, James W.; Schoenfeld, Michael P.; Riley, Daniel P.

2015-01-01

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar

Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase III: Comparison of Theory with Experiment

Science.gov (United States)

Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.

2012-01-01

A proof-of-principle effort to demonstrate a technique by which erosion of the acceleration channel in Hall thrusters of the magnetic-layer type can be eliminated has been completed. The first principles of the technique, now known as "magnetic shielding," were derived based on the findings of numerical simulations in 2-D axisymmetric geometry. The simulations, in turn, guided the modification of an existing 6-kW laboratory Hall thruster. This magnetically shielded (MS) thruster was then built and tested. Because neither theory nor experiment alone can validate fully the first principles of the technique, the objective of the 2-yr effort was twofold: (1) to demonstrate in the laboratory that the erosion rates can be reduced by >order of magnitude, and (2) to demonstrate that the near-wall plasma properties can be altered according to the theoretical predictions. This paper concludes the demonstration of magnetic shielding by reporting on a wide range of comparisons between results from numerical simulations and laboratory diagnostics. Collectively, we find that the comparisons validate the theory. Near the walls of the MS thruster, theory and experiment agree: (1) the plasma potential has been sustained at values near the discharge voltage, and (2) the electron temperature has been lowered by at least 2.5-3 times compared to the unshielded (US) thruster. Also, based on carbon deposition measurements, the erosion rates at the inner and outer walls of the MS thruster are found to be lower by at least 2300 and 1875 times, respectively. Erosion was so low along these walls that the rates were below the resolution of the profilometer. Using a sputtering yield model with an energy threshold of 25 V, the simulations predict a reduction of 600 at the MS inner wall. At the outer wall ion energies are computed to be below 25 V, for which case we set the erosion to zero in the simulations. When a 50-V threshold is used the computed ion energies are below the threshold at both

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

Science.gov (United States)

Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Thomas, Robert; Yim, John; Herman, Daniel; Williams, George; Myers, James; Hofer, Richard;

Thermal Modeling for Pulsed Inductive FRC Plasmoid Thrusters

Science.gov (United States)

Pfaff, Michael

Due to the rising importance of space based infrastructure, long-range robotic space missions, and the need for active attitude control for spacecraft, research into Electric Propulsion is becoming increasingly important. Electric Propulsion (EP) systems utilize electric power to accelerate ions in order to produce thrust. Unlike traditional chemical propulsion, this means that thrust levels are relatively low. The trade-off is that EP thrusters have very high specific impulses (Isp), and can therefore make do with far less onboard propellant than cold gas, monopropellant, or bipropellant engines. As a consequence of the high power levels used to accelerate the ionized propellant, there is a mass and cost penalty in terms of solar panels and a power processing unit. Due to the large power consumption (and waste heat) from electric propulsion thrusters, accurate measurements and predictions of thermal losses are needed. Excessive heating in sensitive locations within a thruster may lead to premature failure of vital components. Between the fixed cost required to purchase these components, as well as the man-hours needed to assemble (or replace) them, attempting to build a high-power thruster without reliable thermal modeling can be expensive. This paper will explain the usage of FEM modeling and experimental tests in characterizing the ElectroMagnetic Plasmoid Thruster (EMPT) and the Electrodeless Lorentz Force (ELF) thruster at the MSNW LLC facility in Redmond, Washington. The EMPT thruster model is validated using an experimental setup, and steady state temperatures are predicted for vacuum conditions. Preliminary analysis of the ELF thruster indicates possible material failure in absence of an active cooling system for driving electronics and for certain power levels.

Hybrid-PIC Computer Simulation of the Plasma and Erosion Processes in Hall Thrusters

Science.gov (United States)

Hofer, Richard R.; Katz, Ira; Mikellides, Ioannis G.; Gamero-Castano, Manuel

2010-01-01

HPHall software simulates and tracks the time-dependent evolution of the plasma and erosion processes in the discharge chamber and near-field plume of Hall thrusters. HPHall is an axisymmetric solver that employs a hybrid fluid/particle-in-cell (Hybrid-PIC) numerical approach. HPHall, originally developed by MIT in 1998, was upgraded to HPHall-2 by the Polytechnic University of Madrid in 2006. The Jet Propulsion Laboratory has continued the development of HPHall-2 through upgrades to the physical models employed in the code, and the addition of entirely new ones. Primary among these are the inclusion of a three-region electron mobility model that more accurately depicts the cross-field electron transport, and the development of an erosion sub-model that allows for the tracking of the erosion of the discharge chamber wall. The code is being developed to provide NASA science missions with a predictive tool of Hall thruster performance and lifetime that can be used to validate Hall thrusters for missions.

Oxygen-Methane Thruster

Science.gov (United States)

Pickens, Tim

2012-01-01

An oxygen-methane thruster was conceived with integrated igniter/injector capable of nominal operation on either gaseous or liquid propellants. The thruster was designed to develop 100 lbf (approximately 445 N) thrust at vacuum conditions and use oxygen and methane as propellants. This continued development included refining the design of the thruster to minimize part count and manufacturing difficulties/cost, refining the modeling tools and capabilities that support system design and analysis, demonstrating the performance of the igniter and full thruster assembly with both gaseous and liquid propellants, and acquiring data from this testing in order to verify the design and operational parameters of the thruster. Thruster testing was conducted with gaseous propellants used for the igniter and thruster. The thruster was demonstrated to work with all types of propellant conditions, and provided the desired performance. Both the thruster and igniter were tested, as well as gaseous propellants, and found to provide the desired performance using the various propellant conditions. The engine also served as an injector testbed for MSFC-designed refractory combustion chambers made of rhenium.

Alkali ion migration between stacked glass plates by corona discharge treatment

Energy Technology Data Exchange (ETDEWEB)

Kawaguchi, Keiga [Research Institute for Electronic Science, Hokkaido University, N20 W10, Kita-ku, Sapporo, Hokkaido 001-0020 (Japan); Suzuki, Toshio [Research Center, Asahi Glass Co., Ltd., 1150 Hazawa-cho, Kanagawa-ku, Yokohama, Kanagawa 221-8755 (Japan); Ikeda, Hiroshi [Art, Science and Technology Center for Cooperative Research, Kyushu University, 6-1 Kasuga-koen, Kasuga, Fukuoka 816-8580 (Japan); Sakai, Daisuke [Department of Electrical and Electronic Engineering, Kitami Institute of Technology, 165 Koen-cho, Kitami, Hokkaido 090-8507 (Japan); Funatsu, Shiro; Uraji, Keiichiro [Production Technology Center, Asahi Glass Co., Ltd., 1-1 Suehiro-cyo, Tsurumiku, Yokohama, Kanagawa 230-0045 (Japan); Yamamoto, Kiyoshi [Research Center, Asahi Glass Co., Ltd., 1150 Hazawa-cho, Kanagawa-ku, Yokohama, Kanagawa 221-8755 (Japan); Harada, Kenji [Department of Computer Science, Kitami Institute of Technology, 165 Koen-cho, Kitami, Hokkaido 090-8507 (Japan); Nishii, Junji, E-mail: nishii@es.hokudai.ac.jp [Research Institute for Electronic Science, Hokkaido University, N20 W10, Kita-ku, Sapporo, Hokkaido 001-0020 (Japan)

2015-05-30

Highlights: • Two stacked glass plates with a 1 mm gap were treated by corona discharge. • Spatial migration of alkali ion over the gap was demonstrated. • Hydrogen gas was necessary for uniform migration. • Surface modification was done with this process without high temperature or vacuum. - Abstract: Corona discharge reflects the spatial migration of alkali ions over a gap between two glass plates. This study examined stacked glass plates containing different alkali ions treated with the corona discharge plasma generated by applied voltage of 4.5 kV at 200 °C. Protons generated at the anode electrode penetrate into the potassium-ion-containing upper glass plate, which is located 5 mm below the anode electrode. Potassium ions intruded into the lower glass plate containing sodium ions placed on the cathode electrode, even over a 1 mm gap separating the plates. Finally, the sodium ion discharged on the cathode electrode. The hydrogen atmosphere was effective at inhibiting the potassium ion reaction with ambient gases during the spatial migration between the two glass plates.

Space Charge Saturated Sheath Regime and Electron Temperature Saturation in Hall Thrusters

International Nuclear Information System (INIS)

Raitses, Y.; Staack, D.; Smirnov, A.; Fisch, N.J.

2005-01-01

Secondary electron emission in Hall thrusters is predicted to lead to space charge saturated wall sheaths resulting in enhanced power losses in the thruster channel. Analysis of experimentally obtained electron-wall collision frequency suggests that the electron temperature saturation, which occurs at high discharge voltages, appears to be caused by a decrease of the Joule heating rather than by the enhancement of the electron energy loss at the walls due to a strong secondary electron emission

Ion clusters, REB, and current sheath characteristics in focused discharges

International Nuclear Information System (INIS)

Bortolotti, A.; Brzosko, J.; DeChiara, P.; Kilic, H.; Mezzetti, F.; Nardi, V.; Powell, C.; Zeng, D.

1990-01-01

Small fluctuations in the current sheath characteristics (peak current density, FWHM of leading sheath, control parameters of sheath internal structure) are linked to wide fluctuations of ion and ion cluster emission from the pinch. Magnetic probe data are used for correlating variations of current sheath parameters with particle emission intensity, Z/M composition, particle energy spectrum. The emission of ion and ion clusters at 90 degrees from the axis of a plasma focus discharge is monitored simultaneously with the 0 degrees emission. The particle energy spectrum is analyzed with a Thomson (parabola) spectrometer (time resolution ∼ 1 nanosec). The cross-sectional structure of the REB at 180 degrees along the discharge axis is monitored via the deposition of collective-field accelerated ions on a target in the REB direction. Etched tracks of ion and ion clusters are in all cases recorded on CR-39 plates. Sharp peaks of the D + -ion spectrum at 90 degrees are found for E > 200 keV/unit charge in all focused discharges. These peaks are due to ion crossing of the azimuthal magnetic field of the pinch region, in a predominant ion cluster structure

Power Dependence of the Electron Mobility Profile in a Hall Thruster

Science.gov (United States)

Jorns, Benjamin A.; Hofery, Richard H.; Mikellides, Ioannis G.

2014-01-01

The electron mobility profile is estimated in a 4.5 kW commercial Hall thruster as a function of discharge power. Internal measurements of plasma potential and electron temperature are made in the thruster channel with a high-speed translating probe. These measurements are presented for a range of throttling conditions from 150 - 400 V and 0.6 - 4.5 kW. The fluid-based solver, Hall2De, is used in conjunction with these internal plasma parameters to estimate the anomalous collision frequency profile at fixed voltage, 300 V, and three power levels. It is found that the anomalous collision frequency profile does not change significantly upstream of the location of the magnetic field peak but that the extent and magnitude of the anomalous collision frequency downstream of the magnetic peak does change with thruster power. These results are discussed in the context of developing phenomenological models for how the collision frequency profile depends on thruster operating conditions.

Discharge behavior of vacuum arc ion source working in pulse mode

International Nuclear Information System (INIS)

Tang Pingying; Dai Jingyi; Tan Xiaohua; Jin Dazhi; Liu Tie; Ding Bonan

2005-01-01

Discharge behavior of the vacuum arc ion source working in pulse mode was investigated using high-speed photography and spectrum diagnosis. The evolvement of cathode spot on hydrogen-impregnated electrode was captured by high-speed photography, and the emission spectra of cathode spot at different pulse currents were analyzed. The experimental results show that in most cases, only one cathode spot can be found in the discharge zone of vacuum arc ion source, and the spot moves a little during the same discharge. Temperature of the cathode spot may rise while the discharge current increases, and ultimately the density of hydrogen ion will be increased. At the same time, sputtering of the electrode is enhanced and the quality of ion plasma will be reduced. (authors)

Study on the hydrogen negative ion in low pressure discharges

International Nuclear Information System (INIS)

Bruneteau, A.M.

1983-07-01

A new use of negative hydrogen ions is the production of intense fast neutral atom beams useful in plasma heating in thermonuclear heating. That is one of the reasons that started this study. The density of negative hydrogen ions in diffusion, and multipole-type low pressure (10 -3 - 10-2 Torr) discharges is deduced from the various formation and destruction processes of the species present in these discharges. The H - ions are essentially produced by dissociative attachment to vibrationally excited molecules and destroyed by processes the relative importance of which is discussed as a function of the discharge parameters. The experimental study of the density of the H - ions, measured by photodetachment, as a function of these parameters, coroborates the theoretical model [fr

Electron and ion kinetics in a micro hollow cathode discharge

Energy Technology Data Exchange (ETDEWEB)

Kim, G J; Iza, F; Lee, J K [Electronics and Electrical Engineering Department, Pohang University of Science and Technology, Pohang, 790-784 (Korea, Republic of)

2006-10-21

Electron and ion kinetics in a micro hollow cathode discharge are investigated by means of two-dimensional axisymmetric particle-in-cell Monte Carlo collision simulations. Argon discharges at 10 and 300 Torr are studied for various driving currents. Electron and ion energy probability functions (IEPF) are shown at various times and locations to study the spatio-temporal behaviour of the discharge. The electron energy probability function (EEPF) evolves from the Druyvesteyn type in the early stages of the discharge into a two (or three) temperature distribution when steady state is reached. In steady state, secondary electrons accelerated across the cathode fall populate the high energy tail of the EEPF while the low energy region is populated by trapped electrons. The IEPF evolves from a Maxwellian in the negative glow (bulk) to a two temperature distribution on the cathode surface. The overpopulation of low energy ions near the cathode surface is attributed to a larger collision cross section for low energy ions and ionization within the cathode fall.

Electron and ion kinetics in a micro hollow cathode discharge

International Nuclear Information System (INIS)

Kim, G J; Iza, F; Lee, J K

2006-01-01

Electron and ion kinetics in a micro hollow cathode discharge are investigated by means of two-dimensional axisymmetric particle-in-cell Monte Carlo collision simulations. Argon discharges at 10 and 300 Torr are studied for various driving currents. Electron and ion energy probability functions (IEPF) are shown at various times and locations to study the spatio-temporal behaviour of the discharge. The electron energy probability function (EEPF) evolves from the Druyvesteyn type in the early stages of the discharge into a two (or three) temperature distribution when steady state is reached. In steady state, secondary electrons accelerated across the cathode fall populate the high energy tail of the EEPF while the low energy region is populated by trapped electrons. The IEPF evolves from a Maxwellian in the negative glow (bulk) to a two temperature distribution on the cathode surface. The overpopulation of low energy ions near the cathode surface is attributed to a larger collision cross section for low energy ions and ionization within the cathode fall

Study of Ion Beam Forming Process in Electric Thruster Using 3D FEM Simulation

Science.gov (United States)

Huang, Tao; Jin, Xiaolin; Hu, Quan; Li, Bin; Yang, Zhonghai

2015-11-01

There are two algorithms to simulate the process of ion beam forming in electric thruster. The one is electrostatic steady state algorithm. Firstly, an assumptive surface, which is enough far from the accelerator grids, launches the ion beam. Then the current density is calculated by theory formula. Secondly these particles are advanced one by one according to the equations of the motions of ions until they are out of the computational region. Thirdly, the electrostatic potential is recalculated and updated by solving Poisson Equation. At the end, the convergence is tested to determine whether the calculation should continue. The entire process will be repeated until the convergence is reached. Another one is time-depended PIC algorithm. In a global time step, we assumed that some new particles would be produced in the simulation domain and its distribution of position and velocity were certain. All of the particles that are still in the system will be advanced every local time steps. Typically, we set the local time step low enough so that the particle needs to be advanced about five times to move the distance of the edge of the element in which the particle is located.

Electronegative Gas Thruster - Direct Thrust Measurement Project

Science.gov (United States)

Dankanich, John (Principal Investigator); Aanesland, Ane; Polzin, Kurt; Walker, Mitchell

2015-01-01

This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct thrust measurement. The initial target application is for Small Satellites, but can be extended to higher power. The Plasma propulsion with Electronegative GASES (PEGASES) concept simplifies ion thruster operation, eliminates a neutralizer requirement and should yield longer life capabilities and lower cost implementation over conventional gridded ion engines. The basic proof-of concept has been demonstrated and matured to TRL 2 over the past several years by researchers at the Laboratoire de Physique des Plasma in France. Due to the low maturity of the innovation, there are currently no domestic investments in electronegative gas thrusters anywhere within NASA, industry or academia. The end product of this Center Innovation Fund (CIF) project will be a validation of the proof-of-concept, maturation to TRL 3 and technology assessment report to summarize the potential for the PEGASES concept to supplant the incumbent technology. Information exchange with the foreign national will be one-way with the exception of the test results. Those test results will first go through a standard public release ITAR/export control review, and the results will be presented in a public technical forum, and the results will be presented in a public technical forum.

High Fidelity Multi-Objective Design Optimization of a Downscaled Cusped Field Thruster

Directory of Open Access Journals (Sweden)

Thomas Fahey

2017-11-01

Full Text Available The Cusped Field Thruster (CFT concept has demonstrated significantly improved performance over the Hall Effect Thruster and the Gridded Ion Thruster; however, little is understood about the complexities of the interactions and interdependencies of the geometrical, magnetic and ion beam properties of the thruster. This study applies an advanced design methodology combining a modified power distribution calculation and evolutionary algorithms assisted by surrogate modeling to a multi-objective design optimization for the performance optimization and characterization of the CFT. Optimization is performed for maximization of performance defined by five design parameters (i.e., anode voltage, anode current, mass flow rate, and magnet radii, simultaneously aiming to maximize three objectives; that is, thrust, efficiency and specific impulse. Statistical methods based on global sensitivity analysis are employed to assess the optimization results in conjunction with surrogate models to identify key design factors with respect to the three design objectives and additional performance measures. The research indicates that the anode current and the Outer Magnet Radius have the greatest effect on the performance parameters. An optimal value for the anode current is determined, and a trend towards maximizing anode potential and mass flow rate is observed.

An experimental investigation of the internal magnetic field topography of an operating Hall thruster

International Nuclear Information System (INIS)

Peterson, Peter Y.; Gallimore, Alec D.; Haas, James M.

2002-01-01

Magnetic field measurements were made in the discharge channel of the 5 kW-class P5 laboratory-model Hall thruster to investigate what effect the Hall current has on the static, applied magnetic field topography. The P5 was operated at 1.6 and 3.0 kW with a discharge voltage of 300 V. A miniature inductive loop probe (B-Dot probe) was employed to measure the radial magnetic field profile inside the discharge channel of the P5 with and without the plasma discharge. These measurements are accomplished with minimal disturbance to thruster operation with the High-speed Axial Reciprocating Probe system. The results of the B-Dot probe measurements indicate a change in the magnetic field topography from that of the vacuum field measurements. The measured magnetic field profiles are then examined to determine the possible nature and source of the difference between the vacuum and plasma magnetic field profiles

Investigation of gas discharge ion sources for on-line mass separation

International Nuclear Information System (INIS)

Kirchner, R.

1976-03-01

The development of efficient gas discharge ion sources with axial beam extraction for on-line mass separation is described. The aim of the investigation was to increase the ion source temperature, the lifetime and the ionisation yield in comparison to present low-pressure are discharge ion sources and to reduce the ion current density from usually 1 to 100 mA/cm 3 . In all ion sources the pressure range below the minimal ignition pressure of the arc discharge was investigated. As a result an ion source was developed which works at small changes in geometry and in electric device of a Nielsen source with high ionization yield (up to 50% for xenon) stabil and without ignition difficulties up to 10 -5 Torr. At a typical pressure of 3 x 10 -5 Torr ion current and ion current density are about 1 μA and 0.1 mA/cm 3 respectively besides high yield and a great emission aperture (diameter 1.2 mm). (orig.) [de

Sources plasma RF magnétisées : applications à la propulsion spatiale

OpenAIRE

Gerst , Jan Dennis

2013-01-01

The PEGASES thruster (Plasma Propulsion with Electronegative Gases) is a novel type of electric thruster for space propulsion. It uses negative and positive ions produced by an inductively coupled radio frequency discharge to create the thrust by electrostatically accelerating the ions through a set of grids. A magnetic filter is used to increase the amount of negative ions in the cavity of the thruster. The PEGASES thruster is not only a source to create a strongly negative ion plasma or eve...

Long Life Cold Cathodes for Hall effect Thrusters, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — An electron source incorporating long life, high current density cold cathodes inside a microchannel plate for use with ion thrusters is proposed. Cathode lifetime...

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

Science.gov (United States)

Garner, Charles E.; Jorns, Benjamin A.; van Derventer, Steven; Hofer, Richard R.; Rickard, Ryan; Liang, Raymond; Delgado, Jorge

2015-01-01

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Beam-plasma discharge in a Kyoto beam-plasma-ion source

International Nuclear Information System (INIS)

Ishikawa, J.; Takagi, T.

1983-01-01

A beam-plasma type ion source employing an original operating principle has been developed by the present authors. The ion source consists of an ion extraction region with an electron gun, a thin long drift tube as the plasma production chamber, and a primary electron beam collector. An electron beam is effectively utilized for the dual purpose of high density plasma production as a result of beam-plasma discharge, and high current ion beam extraction with ion space-charge compensation. A high density plasma of the order of 10 11 --10 13 cm -3 was produced by virtue of the beam-plasma discharge which was caused by the interaction between a space-charge wave on the electron beam and a high frequency plasma wave. The plasma density then produced was 10 2 --10 3 times the density produced only by collisional ionization by the electron beam. In order to obtain a stable beam-plasma discharge, a secondary electron beam emitted from the electron collector should be utilized. The mechanism of the beam-plasma discharge was analyzed by use of a linear theory in the case of the small thermal energy of the electron beam, and by use of a quasilinear theory in the case of the large thermal energy. High current ion beams of more than 0.1 A were extracted even at a low extraction voltage of 1--5 kV

Experimental and theoretical studies of cylindrical Hall thrusters

International Nuclear Information System (INIS)

Smirnov, Artem; Raitses, Yegeny; Fisch, Nathaniel J.

2007-01-01

The Hall thruster is a mature electric propulsion device that holds considerable promise in terms of the propellant saving potential. The annular design of the conventional Hall thruster, however, does not naturally scale to low power. The efficiency tends to be lower and the lifetime issues are more aggravated. Cylindrical geometry Hall thrusters have lower surface-to-volume ratio than conventional thrusters and, thus, seem to be more promising for scaling down. The cylindrical Hall thruster (CHT) is fundamentally different from the conventional design in the way the electrons are confined and the ion space charge is neutralized. The performances of both the large (9-cm channel diameter, 600-1000 W) and miniaturized (2.6-cm channel diameter, 50-300 W) CHTs are comparable with those of the state-of-the-art conventional (annular) design Hall thrusters of similar sizes. A comprehensive experimental and theoretical study of the CHT physics has been conducted, addressing the questions of electron cross-field transport, propellant ionization, plasma-wall interaction, and formation of the electron distribution function. Probe measurements in the harsh plasma environment of the microthruster were performed. Several interesting effects, such as the unusually high ionization efficiency and enhanced electron transport, were observed. Kinetic simulations suggest the existence of the strong fluctuation-enhanced electron diffusion and predict the non-Maxwellian shape of the electron distribution function. Through the acquired understanding of the new physics, ways for further optimization of this means for low-power space propulsion are suggested. Substantial flexibility in the magnetic field configuration of the CHT is the key tool in achieving the high-efficiency operation

Effects of thruster firings on the shuttle's plasma and electric field environment

International Nuclear Information System (INIS)

Machuzak, J.S.; Burke, W.J.; Retterer, J.M.; Hunton, D.E.; Jasperse, J.R.; Smiddy, M.

1993-01-01

Simultaneous plasma and AC/DC electric field measurements taken during the space shuttle mission STS-4 at times of prolonged thruster firings are analyzed and cross correlated. Depending on the orientation of the shuttle's velocity vector to the magnetic field, ion densities and electric field wave spectra were enhanced or decreased. The systematic picture of interactions within the shuttle's plasma/neutral gas environment of Cairns and Gurnett (1991b) is confirmed and extended. Waves are excited by outgassed and thruster-ejected molecules that ionize in close proximity to the shuttle. On time scales significantly less than an ion gyroperiod, the newly created ions act as beams in the background plasma. These beams are sources of VLF waves that propagate near the shuttle and intensify during thruster firings. Plasma density depletions and/or the shuttle's geometry may hinder wave detection in the payload bay. A modified two-stream analysis indicates that beam components propagating at large angles to the magnetic field are unstable to the growth of lower hybrid waves. The beam-excited, lower hybrid waves heat some electrons to sufficient energies to produce impact ionization. Empirical evidence for other wave-growth mechanisms outside the lower-hybrid band is presented. 42 refs., 15 figs., 3 tabs

Combined corona discharge and UV photoionization source for ion mobility spectrometry.

Science.gov (United States)

Bahrami, Hamed; Tabrizchi, Mahmoud

2012-08-15

An ion mobility spectrometer is described which is equipped with two non-radioactive ion sources, namely an atmospheric pressure photoionization and a corona discharge ionization source. The two sources cannot only run individually but are additionally capable of operating simultaneously. For photoionization, a UV lamp was mounted parallel to the axis of the ion mobility cell. The corona discharge electrode was mounted perpendicular to the UV radiation. The total ion current from the photoionization source was verified as a function of lamp current, sample flow rate, and drift field. Simultaneous operation of the two ionization sources was investigated by recording ion mobility spectra of selected samples. The design allows one to observe peaks from either the corona discharge or photoionization individually or simultaneously. This makes it possible to accurately compare peaks in the ion mobility spectra from each individual source. Finally, the instrument's capability for discriminating two peaks appearing in approximately identical drift times using each individual ionization source is demonstrated. Copyright © 2012 Elsevier B.V. All rights reserved.

A Simple Analytical Model for Predicting the Detectable Ion Current in Ion Mobility Spectrometry Using Corona Discharge Ionization Sources

Science.gov (United States)

Kirk, Ansgar Thomas; Kobelt, Tim; Spehlbrink, Hauke; Zimmermann, Stefan

2018-05-01

Corona discharge ionization sources are often used in ion mobility spectrometers (IMS) when a non-radioactive ion source with high ion currents is required. Typically, the corona discharge is followed by a reaction region where analyte ions are formed from the reactant ions. In this work, we present a simple yet sufficiently accurate model for predicting the ion current available at the end of this reaction region when operating at reduced pressure as in High Kinetic Energy Ion Mobility Spectrometers (HiKE-IMS) or most IMS-MS instruments. It yields excellent qualitative agreement with measurement results and is even able to calculate the ion current within an error of 15%. Additional interesting findings of this model are the ion current at the end of the reaction region being independent from the ion current generated by the corona discharge and the ion current in High Kinetic Energy Ion Mobility Spectrometers (HiKE-IMS) growing quadratically when scaling down the length of the reaction region. [Figure not available: see fulltext.

Direction for the Future - Successive Acceleration of Positive and Negative Ions Applied to Space Propulsion

CERN Document Server

Aanesland, A.; Popelier, L.; Chabert, P.

2013-12-16

Electrical space thrusters show important advantages for applications in outer space compared to chemical thrusters, as they allow a longer mission lifetime with lower weight and propellant consumption. Mature technologies on the market today accelerate positive ions to generate thrust. The ion beam is neutralized by electrons downstream, and this need for an additional neutralization system has some drawbacks related to stability, lifetime and total weight and power consumption. Many new concepts, to get rid of the neutralizer, have been proposed, and the PEGASES ion-ion thruster is one of them. This new thruster concept aims at accelerating both positive and negative ions to generate thrust, such that additional neutralization is redundant. This chapter gives an overview of the concept of electric propulsion and the state of the development of this new ion-ion thruster.

Density and velocity measurements of a sheath plasma from MPD thruster

Energy Technology Data Exchange (ETDEWEB)

Ko, J.J.; Cho, T.S.; Choi, M.C.; Choi, E.H.; Cho, G.S.; Uhm, H.S.

1999-07-01

Magnetoplasma is the plasma that the electron and ion orbits are strongly confined by intense magnetic field. Recently, magnetoplasma dynamics (MPD) has been investigated in connection with applications to the rocket thruster in USA, Germany, etc. It can be widely applicable, including modification of satellite position and propulsion of the interplanetary space shuttle. A travel for a long distance journey is possible because a little amount of neutral gases is needed for the plasma source. Besides, this will provide a pollution free engine for future generations. MPD thruster is not a chemical engine. The authors have built a Mather type MPD thruster, which has 1 kV max charging, 10 kA max current flows, and has about 1 ms characteristic operation time. The Paschen curve of this thruster is measured and its minimum breakdown voltage occurs in the pressure range of 0.1 to 1 Torr. Langmuir and double probes are fabricated to diagnose the sheath plasma from the thruster. The temperature and density are calculated to be 2.5 eV and 10{sup 15} cm {sup {minus}3}, respectively, from the probe data. Making use of photo diode, an optical probe is fabricated to measure propagation velocity of the sheath plasma. The sheath plasma from the MPD thruster in the experiment propagates with velocity of 1 cm/{micro}s.

Ion irradiation effects on ionic liquids interfaced with rf discharge plasmas

International Nuclear Information System (INIS)

Baba, K.; Kaneko, T.; Hatakeyama, R.

2007-01-01

The availability of plasma ion irradiation toward a gas-liquid interface is investigated in a rf discharge system incorporating an ionic liquid. The introduction of the ionic liquid to the plasma causes the formation of a sheath electric field on the ionic liquid surface, resulting in the acceleration of the ions to the ionic liquid and the generation of secondary electrons from the ionic liquid by the ion irradiation. These effects are found to advance the discharge process and enhance the plasma production

The nature of the positive ion contribution to a gas discharge

International Nuclear Information System (INIS)

Fletcher, J.; Blevin, H.A.

1980-06-01

The technique for studying swarms of electrons in a gas discharge by observing the photon flux from the discharge developed in the authors's laboratories has been adapted to investigate the role of the secondary mechanisms in hydrogen and nitrogen. The results show that, contrary to previous indications, ion bombardment of the cathode plays only a negligible, if any, part in the low pressure discharge in hydrogen and nitrogen at low E/N while at high E/N only the contribution of the atomic ion is significant

An axially propagating two-stream instability in the Hall thruster plasma

Czech Academy of Sciences Publication Activity Database

Tsikata, S.; Cavalier, Jordan; Héron, A.; Honore, C.; Lemoine, N.; Gresillon, D.; Coulette, D.

2014-01-01

Roč. 21, č. 7 (2014), 072116-072116 ISSN 1070-664X Institutional support: RVO:61389021 Keywords : Collective Thomson scattering * Hall thruster * kinetic theory * electrostatic modes Subject RIV: BL - Plasma and Gas Discharge Physics Impact factor: 2.142, year: 2014 http://dx.doi.org/10.1063/1.4890025

Electrostatic/magnetic ion acceleration through a slowly diverging magnetic nozzle between a ring anode and an on-axis hollow cathode

Directory of Open Access Journals (Sweden)

A. Sasoh

2017-06-01

Full Text Available Ion acceleration through a slowly diverging magnetic nozzle between a ring anode and a hollow cathode set on the axis of symmetry has been realized. Xenon was supplied as the propellant gas from an annular slit along the inner surface of the ring anode so that it was ionized near the anode, and the applied electric potential was efficiently transformed to an ion kinetic energy. As an electrostatic thruster, within the examined operation conditions, the thrust, F, almost scaled with the propellant mass flow rate; the discharge current, Jd, increased with the discharge voltage, Vd. An important characteristic was that the thrust also exhibited electromagnetic acceleration performance, i.e., the so-called “swirl acceleration,” in which F≅JdBRa ∕2, where B and Ra were a magnetic field and an anode inner radius, respectively. Such a unique thruster performance combining both electrostatic and electromagnetic accelerations is expected to be useful as another option for in-space electric propulsion in its broad functional diversity.

1000 Hours of Testing Completed on 10-kW Hall Thruster

Science.gov (United States)

Mason, Lee S.

2001-01-01

Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions.

[Applications of atomic emission spectrum from liquid electrode discharge to metal ion detection].

Science.gov (United States)

Mao, Xiu-Ling; Wu, Jian; Ying, Yi-Bin

2010-02-01

The fast and precise detection of metal ion is an important research project concerning studies in diverse academic fields and different kinds of detecting technologies. In the present paper, the authors review the research on atomic emission spectrum based on liquid electrode discharge and its applications in the detection of metal ion. In the first part of this paper the principles and characteristics of the methods based on electrochemistry and spectroscopy were introduced. The methods of ion-selective electrode (ISE), anodic stripping voltammetry, atomic emission spectrum and atomic absorption spectrum were included in this part and discussed comparatively. Then the principles and characteristics of liquid electrode spectra for metal ion detection were introduced. The mechanism of the plasma production and the characteristics of the plasma spectrum as well as its advantages compared with other methods were discussed. Secondly, the authors divided the discharge system into two types and named them single liquid-electrode discharge and double-liquid electrode respectively, according to the number of the liquid electrode and the configuration of the discharge system, and the development as well as the present research status of each type was illustrated. Then the characteristics and configurations of the discharge systems including ECGD, SCGD, LS-APGD and capillary discharge were discussed in detail as examples of the two types. By taking advantage of the technology of atomic emission spectrum based on liquid electrode discharge, the detecting limit of heavy metals such as copper, mercury and argent as well as active metal ions including sodium, potass and magnesium can achieve microg x L(-1). Finally, the advantages and problems of the liquid-electrode discharge applied in detection of metal ion were discussed. And the applications of the atomic emission spectrum based on liquid electrode discharge were prospected.

Ion Beam Propulsion Study

Science.gov (United States)

2008-01-01

The Ion Beam Propulsion Study was a joint high-level study between the Applied Physics Laboratory operated by NASA and ASRC Aerospace at Kennedy Space Center, Florida, and Berkeley Scientific, Berkeley, California. The results were promising and suggested that work should continue if future funding becomes available. The application of ion thrusters for spacecraft propulsion is limited to quite modest ion sources with similarly modest ion beam parameters because of the mass penalty associated with the ion source and its power supply system. Also, the ion source technology has not been able to provide very high-power ion beams. Small ion beam propulsion systems were used with considerable success. Ion propulsion systems brought into practice use an onboard ion source to form an energetic ion beam, typically Xe+ ions, as the propellant. Such systems were used for steering and correction of telecommunication satellites and as the main thruster for the Deep Space 1 demonstration mission. In recent years, "giant" ion sources were developed for the controlled-fusion research effort worldwide, with beam parameters many orders of magnitude greater than the tiny ones of conventional space thruster application. The advent of such huge ion beam sources and the need for advanced propulsion systems for exploration of the solar system suggest a fresh look at ion beam propulsion, now with the giant fusion sources in mind.

Atomic processes in hydrogen and deuterium negative ion discharges

International Nuclear Information System (INIS)

Hiskes, J.R.

1992-01-01

A knowledge of the atomic processes active in a hydrogen negative ion discharge and their respective rates is an essential component of the interpretation, modeling, and enhancement of negative ion systems. The generation of the cross sections and rate processes appropriate to this problem has been a principal activity at several laboratories. In this paper is discussed those collision processes that are of major importance for the destruction of the vibrationally excited molecules generated in the discharge, processes that are essential to the valuation of the optimization procedure that is to be discussed in this paper

Development of a short pulsed corona discharge ionization source for ion mobility spectrometry

International Nuclear Information System (INIS)

An Yuan; Aliaga-Rossel, R.; Choi, Peter; Gilles, Jean-Paul

2005-01-01

The development of a pulsed corona discharge ionization source and its use in ion mobility spectrometry (IMS) is presented. In a point-plane electrode geometry, an electrical pulse up to 12 kV, 150 ns rise time and 500 ns pulse width was used to generate a corona discharge in air. A single positive high voltage pulse was able to generate about 1.6x10 10 ions at energy consumption of 22 μJ. Since the temporal distribution of ions is in a pulsed form, the possibility of removal the ion gate has been investigated. By purposely arranging the interface between discharge field and drift field, nearly 10 7 positive ions were drawn into the drift region with absence of the ion gate after every single discharge. The positive spectrum of acetone dimer (working at room temperature) was obtained with a resolving power of 20 by using this configuration. The advantages of this new scheme are the low power consumption compared with the dc method as well as the simplicity of the IMS cell structure

Iodine Hall Thruster

Science.gov (United States)

Szabo, James

2015-01-01

Iodine enables dramatic mass and cost savings for lunar and Mars cargo missions, including Earth escape and near-Earth space maneuvers. The demonstrated throttling ability of iodine is important for a singular thruster that might be called upon to propel a spacecraft from Earth to Mars or Venus. The ability to throttle efficiently is even more important for missions beyond Mars. In the Phase I project, Busek Company, Inc., tested an existing Hall thruster, the BHT-8000, on iodine propellant. The thruster was fed by a high-flow iodine feed system and supported by an existing Busek hollow cathode flowing xenon gas. The Phase I propellant feed system was evolved from a previously demonstrated laboratory feed system. Throttling of the thruster between 2 and 11 kW at 200 to 600 V was demonstrated. Testing showed that the efficiency of iodine fueled BHT-8000 is the same as with xenon, with iodine delivering a slightly higher thrust-to-power (T/P) ratio. In Phase II, a complete iodine-fueled system was developed, including the thruster, hollow cathode, and iodine propellant feed system. The nominal power of the Phase II system is 8 kW; however, it can be deeply throttled as well as clustered to much higher power levels. The technology also can be scaled to greater than 100 kW per thruster to support megawatt-class missions. The target thruster efficiency for the full-scale system is 65 percent at high specific impulse (Isp) (approximately 3,000 s) and 60 percent at high thrust (Isp approximately 2,000 s).

Hollow Cathode Assembly Development for the HERMeS Hall Thruster

Science.gov (United States)

Sarver-Verhey, Timothy R.; Kamhawi, Hani; Goebel, Dan M.; Polk, James E.; Peterson, Peter Y.; Robinson, Dale A.

2016-01-01

To support the operation of the HERMeS 12.5 kW Hall Thruster for NASA's Asteroid Redirect Robotic Mission, hollow cathodes using emitters based on barium oxide impregnate and lanthanum hexaboride are being evaluated through wear-testing, performance characterization, plasma modeling, and review of integration requirements. This presentation will present the development approach used to assess the cathode emitter options. A 2,000-hour wear-test of development model Barium Oxide (BaO) hollow cathode is being performed as part of the development plan. Specifically this test is to identify potential impacts cathode emitter life during operation in the HERMeS thruster. The cathode was operated with a magnetic field-equipped anode that simulates the HERMeS hall thruster operating environment. Cathode discharge performance has been stable with the device accumulating 743 hours at the time of this report. Observed voltage changes are attributed to keeper surface condition changes during testing. Cathode behavior during characterization sweeps exhibited stable behavior, including cathode temperature. The details of the cathode assembly operation of the wear-test will be presented.

Experimental investigation of the effects of variable expanding channel on the performance of a low-power cusped field thruster

Directory of Open Access Journals (Sweden)

Hui Liu

2018-04-01

Full Text Available Due to a special magnetic field structure, the multi-cusped field thruster shows advantages of low wall erosion, low noise and high thrust density over a wide range of thrust. In this paper, expanding discharge channels are employed to make up for deficiencies on the range of thrust and plume divergence, which often emerges in conventional straight cylindrical channels. Three thruster geometries are fabricated with different expanding-angle channels, and a group of experiments are carried out to find out their influence on the performance and discharge characteristics of the thruster. A retarding potential analyzer and a Faraday probe are employed to analyze the structures of the plume in these three models. The results show that when the thrusters operate at low mass flow rate, the gradually-expanding channels exhibit lower propellant utilization and lower overall performance by amounts not exceeding 44.8% in ionization rate and 19.5% in anode efficiency, respectively. But the weakening of magnetic field intensity near the exit of expanding channels leads to an extended thrust throttling ability, a smaller plume divergence angle, and a relatively larger stable operating space without mode converting and the consequent performance degradation.

Study on the plasma generation characteristics of an induction-triggered coaxial pulsed plasma thruster

Science.gov (United States)

Weisheng, CUI; Wenzheng, LIU; Jia, TIAN; Xiuyang, CHEN

2018-02-01

At present, spark plugs are used to trigger discharge in pulsed plasma thrusters (PPT), which are known to be life-limiting components due to plasma corrosion and carbon deposition. A strong electric field could be formed in a cathode triple junction (CTJ) to achieve a trigger function under vacuum conditions. We propose an induction-triggered electrode structure on the basis of the CTJ trigger principle. The induction-triggered electrode structure could increase the electric field strength of the CTJ without changing the voltage between electrodes, contributing to a reduction in the electrode breakdown voltage. Additionally, it can maintain the plasma generation effect when the breakdown voltage is reduced in the discharge experiments. The induction-triggered electrode structure could ensure an effective trigger when the ablation distance of Teflon increases, and the magnetic field produced by the discharge current could further improve the plasma density and propagation velocity. The induction-triggered coaxial PPT we propose has a simplified trigger structure, and it is an effective attempt to optimize the micro-satellite thruster.

Ion engine auxiliary propulsion applications and integration study

Science.gov (United States)

Zafran, S. (Editor)

1977-01-01

The benefits derived from application of the 8-cm mercury electron bombardment ion thruster were assessed. Two specific spacecraft missions were studied. A thruster was tested to provide additional needed information on its efflux characteristics and interactive effects. A Users Manual was then prepared describing how to integrate the thruster for auxiliary propulsion on geosynchronous satellites. By incorporating ion engines on an advanced communications mission, the weight available for added payload increases by about 82 kg (181 lb) for a 100 kg (2200 lb) satellite which otherwise uses electrothermal hydrazine. Ion engines can be integrated into a high performance propulsion module that is compatible with the multimission modular spacecraft and can be used for both geosynchronous and low earth orbit applications. The low disturbance torques introduced by the ion engines permit accurate spacecraft pointing with the payload in operation during thrusting periods. The feasibility of using the thruster's neutralizer assembly for neutralization of differentially charged spacecraft surfaces at geosynchronous altitude was demonstrated during the testing program.

Hybrid-Particle-In-Cell Simulation of Backsputtered Carbon Transport in the Near-Field Plume of a Hall Thruster

Science.gov (United States)

Choi, Maria; Yim, John T.; Williams, George J.; Herman, Daniel A.; Gilland, James H.

2018-01-01

Magnetic shielding has eliminated boron nitride erosion as the life limiting mechanism in a Hall thruster but has resulted in erosion of the front magnetic field pole pieces. Recent experiments show that the erosion of graphite pole covers, which are added to protect the magnetic field pole pieces, causes carbon to redeposit on other surfaces, such as boron nitride discharge channel and cathode keeper surfaces. As a part of the risk-reduction activities for Advanced Electric Propulsion System thruster development, this study models transport of backsputtered carbon from the graphite front pole covers and vacuum facility walls. Fluxes, energy distributions, and redeposition rates of backsputtered carbon on the anode, discharge channel, and graphite cathode keeper surfaces are predicted.

Advanced laboratory for testing plasma thrusters and Hall thruster measurement campaign

Directory of Open Access Journals (Sweden)

Szelecka Agnieszka

2016-06-01

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

Diagnostic Setup for Characterization of Near-Anode Processes in Hall Thrusters

International Nuclear Information System (INIS)

Dorf, L.; Raitses, Y.; Fisch, N.J.

2003-01-01

A diagnostic setup for characterization of near-anode processes in Hall-current plasma thrusters consisting of biased and emissive electrostatic probes, high-precision positioning system and low-noise electronic circuitry was developed and tested. Experimental results show that radial probe insertion does not cause perturbations to the discharge and therefore can be used for accurate near-anode measurements

Electronegative Gas Thruster

Science.gov (United States)

Dankanich, John; Polzin, Kurt; Walker, Mitchell

2015-01-01

The project is an international collaboration and academic partnership to mature an innovative electric propulsion thruster concept to Technology Research Level-3 (TRL-3) through direct thrust measurement. The project includes application assessment of the technology ranging from small spacecraft to high power. The Plasma propulsion with Electronegative GASES(PEGASES) basic proof of concept has been matured to TRL-2 by Ane Aanesland of Laboratoire de Physique des Plasma at Ecole Polytechnique. The concept has advantages through eliminating the neutralizer requirement and should yield longer life and lower cost over conventional gridded ion engines. The objective of this research is to validate the proof of concept through the first direct thrust measurements and mature the concept to TRL-3.

Development and Testing of High Current Hollow Cathodes for High Power Hall Thrusters

Science.gov (United States)

Kamhawi, Hani; Van Noord, Jonathan

2012-01-01

NASA's Office of the Chief Technologist In-Space Propulsion project is sponsoring the testing and development of high power Hall thrusters for implementation in NASA missions. As part of the project, NASA Glenn Research Center is developing and testing new high current hollow cathode assemblies that can meet and exceed the required discharge current and life-time requirements of high power Hall thrusters. This paper presents test results of three high current hollow cathode configurations. Test results indicated that two novel emitter configurations were able to attain lower peak emitter temperatures compared to state-of-the-art emitter configurations. One hollow cathode configuration attained a cathode orifice plate tip temperature of 1132 degC at a discharge current of 100 A. More specifically, test and analysis results indicated that a novel emitter configuration had minimal temperature gradient along its length. Future work will include cathode wear tests, and internal emitter temperature and plasma properties measurements along with detailed physics based modeling.

Investigation of plasma potential and pulsed discharge characteristics in enhanced glow discharge plasma immersion ion implantation and deposition

International Nuclear Information System (INIS)

Li Liuhe; Lu Qiuyuan; Fu, Ricky K.Y.; Chu, Paul K.

2009-01-01

Enhanced glow discharge plasma immersion ion implantation and deposition (EGD-PII and D) does not require external plasma sources. In this technique, the plasma is produced by self-glow discharge when a high negative voltage is applied to the sample. The small-area, pointed-shape hollow anode and large area tabular cathode form an electron-focused electric field. Using a special electric field design, the electrons from either the plasma or target (secondary electrons) are focused to a special hollow anode. As a result of the special electron-focusing field, the self-glow discharge process can be enhanced to achieve effective ion implantation into the substrate. In this work, the plasma potential distribution is investigated in details and the possible pulse discharge mechanism is discussed. The unique characteristics of the pulsed plasma and plasma extinction are studied.

Experimental Investigation of a Direct-drive Hall Thruster and Solar Array System at Power Levels up to 10 kW

Science.gov (United States)

Snyder, John S.; Brophy, John R.; Hofer, Richard R.; Goebel, Dan M.; Katz, Ira

2012-01-01

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.

Ion flux nonuniformities in large-area high-frequency capacitive discharges

International Nuclear Information System (INIS)

Perret, A.; Chabert, P.; Booth, J.-P.; Jolly, J.; Guillon, J.; Auvray, Ph.

2003-01-01

Strong nonuniformities of plasma production are expected in capacitive discharges if the excitation wavelength becomes comparable to the reactor size (standing-wave effect) and/or if the plasma skin depth becomes comparable to the plate separation (skin effect) [M. A. Lieberman et al., Plasma Sources Sci. Technol. 11, 283 (2002)]. Ion flux uniformity measurements were carried out in a large-area square (40 cmx40 cm) capacitive discharge driven at frequencies between 13.56 MHz and 81.36 MHz in argon gas at 150 mTorr. At 13.56 MHz, the ion flux was uniform to ±5%. At 60 MHz (and above) and at low rf power, the standing-wave effect was seen (maximum of the ion flux at the center), in good quantitative agreement with theory. At higher rf power, maxima of the ion flux were observed at the edges, due either to the skin effect or to other edge effects

The role of high Rydberg states in the generation of negative ions in negative-ion discharges

International Nuclear Information System (INIS)

Hiskes, J.R.

1995-01-01

The generation of substantial yields of H - ions in a laser excited H 2 gas has been reported by Pinnaduwage and Christoforu. These H - yields have been attributed to (2 + 1) REMP photoexcitation processes leading to dissociative attachment of doubly-excited or superexcited states (SES), or dissociative attachment of high Rydberg product states. The new feature of these experiments is the implied large dissociative attachment rates, of order 10 -6 cm 3 sec -1 , values that are orders-of-magnitude larger than the dissociative attachment of the vibrationally excited levels of the ground electronic state. While these laser excitations are not directly applicable to a hydrogen negative-ion discharge, the implication of large dissociative attachment rates to the high Rydberg states may affect both the total negative-ion density and the interpretation of discharge performance. Within the discharge energetic electrons will collisionally excite the higher Rydberg states, and the relative contribution of the dissociative attachment of these states when compared with the dissociative attachment to the ground state vibrational levels, is the topic of this paper

Design, fabrication and testing of porous tungsten vaporizers for mercury ion thrusters

Science.gov (United States)

Zavesky, R.; Kroeger, E.; Kami, S.

1983-01-01

The dispersions in the characteristics, performance and reliability of vaporizers for early model 30-cm thrusters were investigated. The purpose of the paper is to explore the findings and to discuss the approaches that were taken to reduce the observed dispersion and present the results of a program which validated those approaches. The information that is presented includes porous tungsten materials specifications, a discussion of assembly procedures, and a description of a test program which screens both material and fabrication processes. There are five appendices providing additional detail in the areas of vaporizer contamination, nitrogen flow testing, bubble testing, porosimeter testing, and mercury purity. Four neutralizers, seven cathodes and five main vaporizers were successfully fabricated, tested, and operated on thrusters. Performance data from those devices is presented and indicates extremely repeatable results from using the design and fabrication procedures.

Performance Evaluation of the T6 Ion Engine

Science.gov (United States)

Snyder, John Steven; Goebel, Dan M.; Hofer, Richard R.; Polk, James E.; Wallace, Neil C.; Simpson, Huw

2010-01-01

The T6 ion engine is a 22-cm diameter, 4.5-kW Kaufman-type ion thruster produced by QinetiQ, Ltd., and is baselined for the European Space Agency BepiColombo mission to Mercury and is being qualified under ESA sponsorship for the extended range AlphaBus communications satellite platform. The heritage of the T6 includes the T5 ion thruster now successfully operating on the ESA GOCE spacecraft. As a part of the T6 development program, an engineering model thruster was subjected to a suite of performance tests and plume diagnostics at the Jet Propulsion Laboratory. The engine was mounted on a thrust stand and operated over its nominal throttle range of 2.5 to 4.5 kW. In addition to the typical electrical and flow measurements, an E x B mass analyzer, scanning Faraday probe, thrust vector probe, and several near-field probes were utilized. Thrust, beam divergence, double ion content, and thrust vector movement were all measured at four separate throttle points. The engine performance agreed well with published data on this thruster. At full power the T6 produced 143 mN of thrust at a specific impulse of 4120 seconds and an efficiency of 64%; optimization of the neutralizer for lower flow rates increased the specific impulse to 4300 seconds and the efficiency to nearly 66%. Measured beam divergence was less than, and double ion content was greater than, the ring-cusp-design NSTAR thruster that has flown on NASA missions. The measured thrust vector offset depended slightly on throttle level and was found to increase with time as the thruster approached thermal equilibrium.

Environmental Testing of the NEXT PM1R Ion Engine

Science.gov (United States)

Snyder, John S.; Anderson, John R.; VanNoord, Jonathan L.; Soulas, George C.

2007-01-01

The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 min in each of three axes. Thermal-vacuum testing included three thermal cycles from 120 to 215 C with hot engine re-starts. Thruster performance was nominal throughout the test program, with minor variations in a few engine operating parameters likely caused by facility effects. There were no significant changes in engine performance as characterized by engine operating parameters, ion optics performance measurements, and beam current density measurements, indicating no significant changes to the hardware as a result of the environmental testing. The NEXT PM1R engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. The redesigned cathode heater terminations successfully survived the vibration environments. Based on the results of this test

Electrostatic analyzer for electron and ion energy in glow discharge tube

International Nuclear Information System (INIS)

Bong Kil Yeon.

1984-01-01

The project, the construction and use of an electrostatic energy analyser (Faraday Cup) are described explaining physically its working mechanism. The analyser was used in a glow discharge tube with air and an air-argon mixture. A chapter with the theory of the glow discharge is included. The ion and electron temperatures, the plasma potential and the distribution function for ions and electrons were measured. The electron temperature and plasma potential were also measured using a Langmuir probe and the results show reasonable agreement with the results of the analyser. Good fits of the experimental electron and ion distribution functions were obtained with Maxwellian distributions centered values near the plasma potential. Finally, we discuss the performance of the analyser compared to Langmuir probes. (author) [pt

Ion swarm data for electrical discharge modeling in air and flue gas mixtures

International Nuclear Information System (INIS)

Nelson, D.; Benhenni, M.; Eichwald, O.; Yousfi, M.

2003-01-01

The first step of this work is the determination of the elastic and inelastic ion-molecule collision cross sections for the main ions (N 2 + , O 2 + , CO 2 + , H 2 O + and O - ) usually present either in the air or flue gas discharges. The obtained cross section sets, given for ion kinetic energies not exceeding 100 eV, correspond to the interactions of each ion with its parent molecule (symmetric case) or nonparent molecule (asymmetric case). Then by using these different cross section sets, it is possible to obtain the ion swarm data for the different gas mixtures involving N 2 , CO 2 , H 2 O and O 2 molecules whatever their relative proportions. These ion swarm data are obtained from an optimized Monte Carlo method well adapted for the ion transport in gas mixtures. This also allows us to clearly show that the classical linear approximations usually applied for the ion swarm data in mixtures such as Blanc's law are far to be valid. Then, the ion swarm data are given in three cases of gas mixtures: a dry air (80% N 2 , 20% O 2 ), a ternary gas mixture (82% N 2 , 12% CO 2 , 6% O 2 ) and a typical flue gas (76% N 2 , 12% CO 2 , 6% O 2 , 6% H 2 O). From these reliable ion swarm data, electrical discharge modeling for a wire to plane electrode configuration has been carried out in these three mixtures at the atmospheric pressure for different applied voltages. Under the same discharge conditions, large discrepancies in the streamer formation and propagation have been observed in these three mixture cases. They are due to the deviations existing not only between the different effective electron-molecule ionization rates but also between the ion transport properties mainly because of the presence of a highly polar molecule such as H 2 O. This emphasizes the necessity to properly consider the ion transport in the discharge modeling

Modeling of the near field plume of a Hall thruster

International Nuclear Information System (INIS)

Boyd, Iain D.; Yim, John T.

2004-01-01

In this study, a detailed numerical model is developed to simulate the xenon plasma near-field plume from a Hall thruster. The model uses a detailed fluid model to describe the electrons and a particle-based kinetic approach is used to model the heavy xenon ions and atoms. The detailed model is applied to compute the near field plume of a small, 200 W Hall thruster. Results from the detailed model are compared with the standard modeling approach that employs the Boltzmann model. The usefulness of the model detailed is assessed through direct comparisons with a number of different measured data sets. The comparisons illustrate that the detailed model accurately predicts a number of features of the measured data not captured by the simpler Boltzmann approach

Thermo-structural modelling of a plasma discharge tube for electric propulsion

International Nuclear Information System (INIS)

Faoite, D. de; Browne, D.J.; Del Valle Gamboa, J.I.; Stanton, K.T.

2016-01-01

Highlights: • Thermo-structural analyses were performed for an electric propulsion space thruster. • Thermal stresses arise primarily from mismatches in thermal expansion coefficients. • Aluminium nitride is a suitable material for a plasma containment tube. • A design is presented allowing a thruster to operate at a power of at least 250 kW. - Abstract: Potential thermal management strategies for the plasma generation section of a VASIMR"® high-power electric propulsion space thruster are assessed. The plasma is generated in a discharge tube using helicon waves. The plasma generation process causes a significant thermal load on the plasma discharge tube and on neighbouring components, caused by cross-field particle diffusion and UV radiation. Four potential cooling system design strategies are assessed to deal with this thermal load. Four polycrystalline ceramics are evaluated for use as the plasma discharge tube material: alumina, aluminium nitride, beryllia, and silicon nitride. A finite element analysis (FEA) method was used to model the steady-state temperature and stress fields resulting from the plasma heat flux. Of the four materials assessed, aluminium nitride would result in the lowest plasma discharge tube temperatures and stresses. It was found that a design consisting of a monolithic ceramic plasma containment tube fabricated from aluminium nitride would be capable of operating up to a power level of at least 250 kW.

Diagnostics of discharge channels for neutralized chamber transport in heavy ion fusion

International Nuclear Information System (INIS)

Niemann, C.; Penache, D.; Tauschwitz, A.; Rosmej, F.B.; Neff, S.; Birkner, R.; Constantin, C.; Knobloch, R.; Presura, R.; Yu, S.S.; Sharp, W.M.; Ponce, D.M.; Hoffmann, D.H.H.

2002-01-01

The final beam transport in the reactor chamber for heavy ion fusion in preformed plasma channels offers many attractive advantages compared to other transport modes. In the past few years, experiments at the Gesellschaft fuer Schwerionenforschung (GSI) accelerator facility have addressed the creation and investigation of discharge plasmas, designed for the transport of intense ion beams. Stable, self-standing channels of 50 cm length with currents up to 55 kA were initiated in low-pressure ammonia gas by a CO 2 -laser pulse along the channel axis before the discharge is triggered. The channels were characterized by several plasma diagnostics including interferometry and spectroscopy. We also present first experiments on laser-guided intersecting discharges

Numerical study on the electron—wall interaction in a Hall thruster with segmented electrodes placed at the channel exit

International Nuclear Information System (INIS)

Qing Shao-Wei; E Peng; Xu Dian-Guo; Duan Ping

2013-01-01

Electron—wall interaction is always recognized as an important physical problem because of its remarkable influences on thruster discharge and performance. Based on existing theories, an electrode is predicted to weaken electron—wall interaction due to its low secondary electron emission characteristic. In this paper, the electron—wall interaction in an Aton-type Hall thruster with low-emissive electrodes placed near the exit of discharge channel is studied by a fully kinetic particle-in-cell method. The results show that the electron—wall interaction in the region of segmented electrode is indeed weakened, but it is significantly enhanced in the remaining region of discharge channel. It is mainly caused by electrode conductive property which makes equipotential lines convex toward channel exit and even parallel to wall surface in near-wall region; this convex equipotential configuration results in significant physical effects such as repelling electrons, which causes the electrons to move toward the channel center, and the electrons emitted from electrodes to be remarkably accelerated, thereby increasing electron temperature in the discharge channel, etc. Furthermore, the results also indicate that the discharge current in the segmented electrode case is larger than in the non-segmented electrode case, which is qualitatively in accordance with previous experimental results. (physics of gases, plasmas, and electric discharges)

Detection of negative ions in glow discharge mass spectrometry for analysis of solid specimens

DEFF Research Database (Denmark)

Canulescu, Stela; Molchan, Igor S.; Tauziede, C.

2010-01-01

A new method is presented for elemental and molecular analysis of halogen-containing samples by glow discharge time-of-flight mass spectrometry, consisting of detection of negative ions from a pulsed RF glow discharge in argon. Analyte signals are mainly extracted from the afterglow regime...... be used to study the distribution of a tantalum fluoride layer within the anodized tantala layer. Further, comparison is made with data obtained using glow-discharge optical emission spectroscopy, where elemental fluorine can only be detected using a neon plasma. The ionization mechanisms responsible...... for the formation of negative ions in glow discharge time-of-flight mass spectrometry are briefly discussed....

Electron temperature measurement in Maxwellian non-isothermal beam plasma of an ion thruster

International Nuclear Information System (INIS)

Zhang, Zun; Tang, Haibin; Kong, Mengdi; Zhang, Zhe; Ren, Junxue

2015-01-01

Published electron temperature profiles of the beam plasma from ion thrusters reveal many divergences both in magnitude and radial variation. In order to know exactly the radial distributions of electron temperature and understand the beam plasma characteristics, we applied five different experimental approaches to measure the spatial profiles of electron temperature and compared the agreement and disagreement of the electron temperature profiles obtained from these techniques. Experimental results show that the triple Langmuir probe and adiabatic poly-tropic law methods could provide more accurate space-resolved electron temperature of the beam plasma than other techniques. Radial electron temperature profiles indicate that the electrons in the beam plasma are non-isothermal, which is supported by a radial decrease (∼2 eV) of electron temperature as the plume plasma expands outward. Therefore, the adiabatic “poly-tropic law” is more appropriate than the isothermal “barometric law” to be used in electron temperature calculations. Moreover, the calculation results show that the electron temperature profiles derived from the “poly-tropic law” are in better agreement with the experimental data when the specific heat ratio (γ) lies in the range of 1.2-1.4 instead of 5/3

RHETT2/EPDM Hall Thruster Propulsion System Electromagnetic Compatibility Evaluation

Science.gov (United States)

Sarmiento, Charles J.; Sankovic, John M.; Freitas, Joseph; Lynn, Peter R.

1997-01-01

Electromagnetic compatibility measurements were obtained as part of the Electric Propulsion Demonstration Module (EPDM) flight qualification program. Tests were conducted on a Hall thruster system operating at a nominal 66O W discharge power. Measurements of conducted and radiated susceptibility and emissions were obtained and referenced to MEL-STD-461 C. The power processor showed some conducted susceptibility below 4 kHz for the magnet current and discharge voltage. Radiated susceptibility testing yielded a null result. Conducted emissions showed slight violations of the specified limit for MIL-461C CE03. Radiated emissions exceeded the RE02 standard at low frequencies, below 300 MHz, by up to 40 dB RV/m/MHz.

A cold cathode of a gas-discharge electron-ion gun

International Nuclear Information System (INIS)

1974-01-01

A cold cathode of a gas-discharge electron-ion gun is constructed in order to continuously replace the eroded material by feeding a wire or a set of coaxial cylinders in the spot where the ions hit the cathode. In this way, the form of the cathode and the electric-field configuration is preserved which guarantees the conservation of a sharp narrow electron beam profile

Volume generation of negative ions in high density hydrogen discharges

International Nuclear Information System (INIS)

Hiskes, J.R.; Karo, A.M.

1983-01-01

A parametric survey is made of a high-density tandem two-chamber hydrogen negative ion system. The optimum extracted negative ion current densities are sensitive to the atom concentration in the discharge and to the system scale length. For scale lengths ranging from 10 cm to 0.1 cm optimum current densities range from of order 1 to 100 mA cm -2 , respectively

Formation of ECR Plasma in a Dielectric Plasma Guide under Self-Excitation of a Standing Ion-Acoustic Wave

Science.gov (United States)

Balmashnov, A. A.; Kalashnikov, A. V.; Kalashnikov, V. V.; Stepina, S. P.; Umnov, A. M.

2018-01-01

The formation of a spatially localized plasma with a high brightness has been experimentally observed in a dielectric plasma guide under the electron cyclotron resonance discharge at the excitation of a standing ion-acoustic wave. The results obtained show the possibility of designing compact high-intensity radiation sources with a spectrum determined by the working gas or gas mixture type, high-intensity chemically active particle flow sources, and plasma thrusters for correcting orbits of light spacecraft.

Neutralization of an ion beam from the end-Hall ion source by a plasma electron source based on a discharge in crossed E × H fields

Science.gov (United States)

Dostanko, A. P.; Golosov, D. A.

2009-10-01

The possibility of using a plasma electron source (PES) with a discharge in crossed E × H field for compensating the ion beam from an end-Hall ion source (EHIS) is analyzed. The PES used as a neutralizer is mounted in the immediate vicinity of the EHIS ion generation and acceleration region at 90° to the source axis. The behavior of the discharge and emission parameters of the EHIS is determined for operation with a filament neutralizer and a plasma electron source. It is found that the maximal discharge current from the ion source attains a value of 3.8 A for operation with a PES and 4 A for operation with a filament compensator. It is established that the maximal discharge current for the ion source strongly depends on the working gas flow rate for low flow rates (up to 10 ml/min) in the EHIS; for higher flow rates, the maximum discharge current in the EHIS depends only on the emissivity of the PES. Analysis of the emission parameters of EHISs with filament and plasma neutralizers shows that the ion beam current and the ion current density distribution profile are independent of the type of the electron source and the ion current density can be as high as 0.2 mA/cm2 at a distance of 25 cm from the EHIS anode. The balance of currents in the ion source-electron source system is considered on the basis of analysis of operation of EHISs with various sources of electrons. It is concluded that the neutralization current required for operation of an ion source in the discharge compensation mode must be equal to or larger than the discharge current of the ion source. The use of PES for compensating the ion beam from an end-Hall ion source proved to be effective in processes of ion-assisted deposition of thin films using reactive gases like O2 or N2. The application of the PES technique makes it possible to increase the lifetime of the ion-assisted deposition system by an order of magnitude (the lifetime with a Ti cathode is at least 60 h and is limited by the

Brine reuse in ion-exchange softening: salt discharge, hardness leakage, and capacity tradeoffs.

Science.gov (United States)

Flodman, Hunter R; Dvorak, Bruce I

2012-06-01

Ion-exchange water softening results in the discharge of excess sodium chloride to the aquatic environment during the regeneration cycle. In order to reduce sodium chloride use and subsequent discharge from ion-exchange processes, either brine reclaim operations can be implemented or salt application during regeneration can be reduced. Both result in tradeoffs related to loss of bed volumes treated per cycle and increased hardness leakage. An experimentally validated model was used to compare concurrent water softening operations at various salt application quantities with and without the direct reuse of waste brine for treated tap water of typical midwestern water quality. Both approaches were able to reduce salt use and subsequent discharge. Reducing salt use and discharge by lowering the salt application rate during regeneration consequently increased hardness leakage and decreased treatment capacity. Single or two tank brine recycling systems are capable of reducing salt use and discharge without increasing hardness leakage, although treatment capacity is reduced.

Oxygen-Methane Thruster, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Orion Propulsion, Inc. proposes to develop an Oxygen and Methane RCS Thruster to advance the technology of alternate fuels. A successful Oxygen/CH4 RCS Thruster will...

Numerical investigation of thermal behaviors in lithium-ion battery stack discharge

International Nuclear Information System (INIS)

Liu, Rui; Chen, Jixin; Xun, Jingzhi; Jiao, Kui; Du, Qing

2014-01-01

Highlights: • The thermal behaviors of a Li-ion battery stack have been investigated by modeling. • Parametric studies have been performed focusing on three different cooling materials. • Effects of discharge rate, ambient temperature and Reynolds number are examined. • General guidelines are proposed for the thermal management of a Li-ion battery stack. - Abstract: Thermal management is critically important to maintain the performance and prolong the lifetime of a lithium-ion (Li-ion) battery. In this paper, a two-dimensional and transient model has been developed for the thermal management of a 20-flat-plate-battery stack, followed by comprehensive numerical simulations to study the influences of ambient temperature, Reynolds number, and discharge rate on the temperature distribution in the stack with different cooling materials. The simulation results indicate that liquid cooling is generally more effective in reducing temperature compared to phase-change material, while the latter can lead to more homogeneous temperature distribution. Fast and deep discharge should be avoided, which generally yields high temperature beyond the acceptable range regardless of cooling materials. At low or even subzero ambient temperatures, air cooling is preferred over liquid cooling because heat needs to be retained rather than removed. Such difference becomes small when the ambient temperature increases to a mild level. The effects of Reynolds number are apparent in liquid cooling but negligible in air cooling. Choosing appropriate cooling material and strategy is particularly important in low ambient temperature and fast discharge cases. These findings improve the understanding of battery stack thermal behaviors and provide the general guidelines for thermal management system. The present model can also be used in developing control system to optimize battery stack thermal behaviors

Analysis of the ion energy transport in ohmic discharges in the ASDEX tokamak

International Nuclear Information System (INIS)

Simmet, E.E.; Fahrbach, H.U.; Herrmann, W.; Stroth, U.

1996-10-01

An analysis of the local ion energy transport is performed for more than one hundred well documented ohmic ASDEX discharges. These are characterized by three different confinement regimes: the linear ohmic confinement (LOC), the saturated ohmic confinement (SOC) and the improved ohmic confinement (IOC). All three are covered by this study. To identify the most important local transport mechanism of the ion heat, the ion power balance equation is analyzed. Two methods are used: straightforward calculation with experimental data only, and a comparison of measured and calculated profiles of the ion temperature and the ion heat conductivity, respectively. A discussion of the power balance shows that conductive losses dominate the ion energy transport in all ohmic discharges of ASDEX. Only inside the q=1-surface losses due to sawtooth activity play a role, while at the edge convective fluxes and CX-losses influence the ion energy transport. Both methods lead to the result that both the ion temperature and the ion heat conductivity are consistent with predictions of the neoclassical theory. Enhanced heat losses as suggested by theories eg. on the basis of η i modes can be excluded. (orig.)

Non-Maxwellian electron energy probability functions in the plume of a SPT-100 Hall thruster

Science.gov (United States)

Giono, G.; Gudmundsson, J. T.; Ivchenko, N.; Mazouffre, S.; Dannenmayer, K.; Loubère, D.; Popelier, L.; Merino, M.; Olentšenko, G.

2018-01-01

We present measurements of the electron density, the effective electron temperature, the plasma potential, and the electron energy probability function (EEPF) in the plume of a 1.5 kW-class SPT-100 Hall thruster, derived from cylindrical Langmuir probe measurements. The measurements were taken on the plume axis at distances between 550 and 1550 mm from the thruster exit plane, and at different angles from the plume axis at 550 mm for three operating points of the thruster, characterized by different discharge voltages and mass flow rates. The bulk of the electron population can be approximated as a Maxwellian distribution, but the measured distributions were seen to decline faster at higher energy. The measured EEPFs were best modelled with a general EEPF with an exponent α between 1.2 and 1.5, and their axial and angular characteristics were studied for the different operating points of the thruster. As a result, the exponent α from the fitted distribution was seen to be almost constant as a function of the axial distance along the plume, as well as across the angles. However, the exponent α was seen to be affected by the mass flow rate, suggesting a possible relationship with the collision rate, especially close to the thruster exit. The ratio of the specific heats, the γ factor, between the measured plasma parameters was found to be lower than the adiabatic value of 5/3 for each of the thruster settings, indicating the existence of non-trivial kinetic heat fluxes in the near collisionless plume. These results are intended to be used as input and/or testing properties for plume expansion models in further work.

Simulation of electrochemical behavior in Lithium ion battery during discharge process.

Science.gov (United States)

Chen, Yong; Huo, Weiwei; Lin, Muyi; Zhao, Li

2018-01-01

An electrochemical Lithium ion battery model was built taking into account the electrochemical reactions. The polarization was divided into parts which were related to the solid phase and the electrolyte mass transport of species, and the electrochemical reactions. The influence factors on battery polarization were studied, including the active material particle radius and the electrolyte salt concentration. The results showed that diffusion polarization exist in the positive and negative electrodes, and diffusion polarization increase with the conducting of the discharge process. The physicochemical parameters of the Lithium ion battery had the huge effect on cell voltage via polarization. The simulation data show that the polarization voltage has close relationship with active material particle size, discharging rate and ambient temperature.

Experimental study of effect of magnetic field on anode temperature distribution in an ATON-type Hall thruster

Science.gov (United States)

Liu, Jinwen; Li, Hong; Mao, Wei; Ding, Yongjie; Wei, Liqiu; Li, Jianzhi; Yu, Daren; Wang, Xiaogang

2018-05-01

The energy deposition caused by the absorption of electrons by the anode is an important cause of power loss in a Hall thruster. The resulting anode heating is dangerous, as it can potentially reduce the thruster lifetime. In this study, by considering the ring shape of the anode of an ATON-type Hall thruster, the effects of the magnetic field strength and gradient on the anode ring temperature distribution are studied via experimental measurement. The results show that the temperature distribution is not affected by changes in the magnetic field strength and that the position of the peak temperature is essentially unchanged; however, the overall temperature does not change monotonically with the increase of the magnetic field strength and is positively correlated with the change in the discharge current. Moreover, as the magnetic field gradient increases, the position of the peak temperature gradually moves toward the channel exit and the temperature tends to decrease as a whole, regardless of the discharge current magnitude; in any case, the position of the peak temperature corresponds exactly to the intersection of the magnetic field cusp with the anode ring. Further theoretical analysis shows that the electrons, coming from the ionization region, travel along two characteristic paths to reach the anode under the guidance of the cusped magnetic field configuration. The change of the magnetic field strength or gradient changes the transfer of momentum and energy of the electrons in these two paths, which is the main reason for the changes in the temperature and distribution. This study is instructive for matching the design of the ring-shaped anode and the cusp magnetic field of an ATON-type Hall thruster.

Scale Model Thruster Acoustic Measurement Results

Science.gov (United States)

Vargas, Magda; Kenny, R. Jeremy

2013-01-01

The Space Launch System (SLS) Scale Model Acoustic Test (SMAT) is a 5% scale representation of the SLS vehicle, mobile launcher, tower, and launch pad trench. The SLS launch propulsion system will be comprised of the Rocket Assisted Take-Off (RATO) motors representing the solid boosters and 4 Gas Hydrogen (GH2) thrusters representing the core engines. The GH2 thrusters were tested in a horizontal configuration in order to characterize their performance. In Phase 1, a single thruster was fired to determine the engine performance parameters necessary for scaling a single engine. A cluster configuration, consisting of the 4 thrusters, was tested in Phase 2 to integrate the system and determine their combined performance. Acoustic and overpressure data was collected during both test phases in order to characterize the system's acoustic performance. The results from the single thruster and 4- thuster system are discussed and compared.

Laser Induced Fluorescence Measurements in a Hall Thruster Plume as a Function of Background Pressure

Science.gov (United States)

Spektor, R.; Tighe, W. G.; Kamhawi, H.

2016-01-01

A set of Laser Induced Fluorescence (LIF) measurements in the near-field region of the NASA- 173M Hall thruster plume is presented at four background pressure conditions varying from 9.4 x 10(exp -6) torr to 3.3 x 10(exp -5) torr. The xenon ion velocity distribution function was measured simultaneously along the axial and radial directions. An ultimate exhaust velocity of 19.6+/-0.25 km/s achieved at a distance of 20 mm was measured, and that value was not sensitive to pressure. On the other hand, the ion axial velocity at the thruster exit was strongly influenced by pressure, indicating that the accelerating electric field moved inward with increased pressure. The shift in electric field corresponded to an increase in measured thrust. Pressure had a minor effect on the radial component of ion velocity, mainly affecting ions exiting close to the channel inner wall. At that radial location the radial component of ion velocity was approximately 1000 m/s greater at the lowest pressure than at the highest pressure. A reduction of the inner magnet coil current by 0.6 A resulted in a lower axial ion velocity at the channel exit while the radial component of ion velocity at the channel inner wall location increased by 1300 m/s, and at the channel outer wall location the radial ion velocity remained unaffected. The ultimate exhaust velocity was not significantly affected by the inner magnet current.

Q-Thruster Breadboard Campaign Project

Science.gov (United States)

White, Harold

2014-01-01

Dr. Harold "Sonny" White has developed the physics theory basis for utilizing the quantum vacuum to produce thrust. The engineering implementation of the theory is known as Q-thrusters. During FY13, three test campaigns were conducted that conclusively demonstrated tangible evidence of Q-thruster physics with measurable thrust bringing the TRL up from TRL 2 to early TRL 3. This project will continue with the development of the technology to a breadboard level by leveraging the most recent NASA/industry test hardware. This project will replace the manual tuning process used in the 2013 test campaign with an automated Radio Frequency (RF) Phase Lock Loop system (precursor to flight-like implementation), and will redesign the signal ports to minimize RF leakage (improves efficiency). This project will build on the 2013 test campaign using the above improvements on the test implementation to get ready for subsequent Independent Verification and Validation testing at Glenn Research Center (GRC) and Jet Propulsion Laboratory (JPL) in FY 2015. Q-thruster technology has a much higher thrust to power than current forms of electric propulsion (7x Hall thrusters), and can significantly reduce the total power required for either Solar Electric Propulsion (SEP) or Nuclear Electric Propulsion (NEP). Also, due to the high thrust and high specific impulse, Q-thruster technology will greatly relax the specific mass requirements for in-space nuclear reactor systems. Q-thrusters can reduce transit times for a power-constrained architecture.

Effect of plasma distribution on propulsion performance in electrodeless plasma thrusters

Science.gov (United States)

Takao, Yoshinori; Takase, Kazuki; Takahashi, Kazunori

2016-09-01

A helicon plasma thruster consisting of a helicon plasma source and a magnetic nozzle is one of the candidates for long-lifetime thrusters because no electrodes are employed to generate or accelerate plasma. A recent experiment, however, detected the non-negligible axial momentum lost to the lateral wall boundary, which degrades thruster performance, when the source was operated with highly ionized gases. To investigate this mechanism, we have conducted two-dimensional axisymmetric particle-in-cell (PIC) simulations with the neutral distribution obtained by Direct Simulation Monte Carlo (DSMC) method. The numerical results have indicated that the axially asymmetric profiles of the plasma density and potential are obtained when the strong decay of neutrals occurs at the source downstream. This asymmetric potential profile leads to the accelerated ion towards the lateral wall, leading to the non-negligible net axial force in the opposite direction of the thrust. Hence, to reduce this asymmetric profile by increasing the neutral density at downstream and/or by confining plasma with external magnetic field would result in improvement of the propulsion performance. These effects are also analyzed by PIC/DSMC simulations.

Current limitation by an electric double layer in ion laser discharges

International Nuclear Information System (INIS)

Torven, S.

1977-12-01

A theory for current limitation in ion laser discharges is investigated. The basic mechanism considered is saturation of the positive ion flux at an electric double layer by the limited flux of neutral atoms. The result is compared with a recently published synthesis of a large number of experimental data which agree well with those predicted by the double layer model

Numerical investigation of a Hall thruster plasma

International Nuclear Information System (INIS)

Roy, Subrata; Pandey, B.P.

2002-01-01

The dynamics of the Hall thruster is investigated numerically in the framework of a one-dimensional, multifluid macroscopic description of a partially ionized xenon plasma using finite element formulation. The model includes neutral dynamics, inelastic processes, and plasma-wall interaction. Owing to disparate temporal scales, ions and neutrals have been described by set of time-dependent equations, while electrons are considered in steady state. Based on the experimental observations, a third order polynomial in electron temperature is used to calculate ionization rate. The results show that in the acceleration channel the increase in the ion number density is related to the decrease in the neutral number density. The electron and ion velocity profiles are consistent with the imposed electric field. The electron temperature remains uniform for nearly two-thirds of the channel; then sharply increases to a peak before dropping slightly at the exit. This is consistent with the predicted electron gyration velocity distribution

Charge–discharge properties of tin dioxide for sodium-ion battery

Energy Technology Data Exchange (ETDEWEB)

Park, Jinsoo [Department of Materials and Energy Engineering, Kyungwoon University, 730 Gangdong-ro, Sandong-meon, Gumi-si, Gyeongbuk 730-739 (Korea, Republic of); Park, Jin-Woo; Han, Jeong-Hui [School of Materials Science and Engineering, RIGET, Gyeongsang National University, 900 Gajwa-dong, Jinju 660-701 (Korea, Republic of); Lee, Sang-Won [R and D 2 Team, COSMO AM and T CO., LTD., 315 Mokhaeng-dong, Chungju (Korea, Republic of); Lee, Ki-Young [Jeonyoung ECP, 637-1, Sunggok-dong, Danwon-gu, Ansan cilt, Kyunggi-do (Korea, Republic of); Ryu, Ho-Suk; Kim, Ki-Won [School of Materials Science and Engineering, RIGET, Gyeongsang National University, 900 Gajwa-dong, Jinju 660-701 (Korea, Republic of); Wang, Guoxiu [School of Chemistry and Forensic Science, University of Technology Sydney, Sydney, NSW 2007 (Australia); Ahn, Jou-Hyeon [Department of Chemical and Biological Engineering, Gyeongsang National University, 900 Gajwa-dong, Jinju 660-701 (Korea, Republic of); Ahn, Hyo-Jun, E-mail: ahj@gnu.ac.kr [School of Materials Science and Engineering, RIGET, Gyeongsang National University, 900 Gajwa-dong, Jinju 660-701 (Korea, Republic of)

2014-10-15

Highlights: • The electrochemical reaction of SnO2 as an anode for Na-ion batteries was studied. • The SnO2 electrode delivered the initial discharge capacity of 747 mAh/g. • Alarge irreversible capacity (597 mAh/g)was observedin the first cycle. • The in-plain crack in the electrode caused the incompletereduction of SnO{sub 2}. - Abstract: Tin dioxide was investigated as an anode material for sodium-ion batteries. The Na/SnO{sub 2} cell delivered a first discharge capacity of 747 mAh/g, but the first charge capacity was 150 mAh/g. The irreversible capacity in the first cycle was examined through characterization by X-ray diffraction and scanning electron microscopy. X-ray diffraction analysis revealed that the SnO{sub 2} active material was not reduced fully to metallic Sn. Furrows and wrinkles were formed on the electrode surface owing to the volumetric expansion upon first discharge, which led to a deterioration of the electrode structure and a loss of electrical contact between the active materials. The analysis is summarized in the schematic drawing.

Ion transport analysis of a high beta-poloidal JT-60U discharge

International Nuclear Information System (INIS)

Horton, W.; Tajima, T.; Dong, J.-Q.; Kim, J.-Y.; Kishimoto, Y.

1997-01-01

The high beta-poloidal discharge number 17110 in JT-60U (JT-60 Team, IAEA, Vienna, 1993) that developes an internal transport barrier is analysed for the transport of ion energy and momentum. First, the classical ion temperature gradient stability properties are calculated in the absence of sheared plasma flows to establish the L-mode transport level prior to the emergence of the transport barrier. Then the evolving toroidal and poloidal velocity profiles reported by Koide et al (1994 Phys. Rev. Lett. 72 3662) are used to show how the sheared mass flows control the stability and transport. Coupled energy-momentum transport equations predict the creation of a transport barrier. The balance of the steep ion temperature gradient against the magnetic shear and sheared mass flow is calculated for the profiles in the 17110 discharge. (Author)

High-Power Krypton Hall Thruster Technology Being Developed for Nuclear-Powered Applications

Science.gov (United States)

Jacobson, David T.; Manzella, David H.

2004-01-01

The NASA Glenn Research Center has been performing research and development of moderate specific impulse, xenon-fueled, high-power Hall thrusters for potential solar electric propulsion applications. These applications include Mars missions, reusable tugs for low-Earth-orbit to geosynchronous-Earth-orbit transportation, and missions that require transportation to libration points. This research and development effort resulted in the design and fabrication of the NASA-457M Hall thruster that has been tested at input powers up to 95 kW. During project year 2003, NASA established Project Prometheus to develop technology in the areas of nuclear power and propulsion, which are enabling for deep-space science missions. One of the Project-Prometheus-sponsored Nuclear Propulsion Research tasks is to investigate alternate propellants for high-power Hall thruster electric propulsion. The motivation for alternate propellants includes the disadvantageous cost and availability of xenon propellant for extremely large scale, xenon-fueled propulsion systems and the potential system performance benefits of using alternate propellants. The alternate propellant krypton was investigated because of its low cost relative to xenon. Krypton propellant also has potential performance benefits for deep-space missions because the theoretical specific impulse for a given voltage is 20 percent higher than for xenon because of krypton's lower molecular weight. During project year 2003, the performance of the high-power NASA-457M Hall thruster was measured using krypton as the propellant at power levels ranging from 6.4 to 72.5 kW. The thrust produced ranged from 0.3 to 2.5 N at a discharge specific impulse up to 4500 sec.

NASA HERMeS Hall Thruster Electrical Configuration Characterization

Science.gov (United States)

Peterson, Peter; Kamhawi, Hani; Huang, Wensheng; Yim, John; Herman, Daniel; Williams, George; Gilland, James; Hofer, Richard

2016-01-01

NASAs Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This presentation will cover the electrical configuration testing of the TDU-1 HERMeS Hall thruster in NASA Glenn Research Centers Vacuum Facility 5. The three electrical configurations examined are the thruster body tied to facility ground, thruster floating, and finally the thruster body electrically tied to cathode common. The TDU-1 HERMeS was configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

Anomalous electron transport in Hall-effect thrusters: Comparison between quasi-linear kinetic theory and particle-in-cell simulations

Science.gov (United States)

Lafleur, T.; Martorelli, R.; Chabert, P.; Bourdon, A.

2018-06-01

Kinetic drift instabilities have been implicated as a possible mechanism leading to anomalous electron cross-field transport in E × B discharges, such as Hall-effect thrusters. Such instabilities, which are driven by the large disparity in electron and ion drift velocities, present a significant challenge to modelling efforts without resorting to time-consuming particle-in-cell (PIC) simulations. Here, we test aspects of quasi-linear kinetic theory with 2D PIC simulations with the aim of developing a self-consistent treatment of these instabilities. The specific quantities of interest are the instability growth rate (which determines the spatial and temporal evolution of the instability amplitude), and the instability-enhanced electron-ion friction force (which leads to "anomalous" electron transport). By using the self-consistently obtained electron distribution functions from the PIC simulations (which are in general non-Maxwellian), we find that the predictions of the quasi-linear kinetic theory are in good agreement with the simulation results. By contrast, the use of Maxwellian distributions leads to a growth rate and electron-ion friction force that is around 2-4 times higher, and consequently significantly overestimates the electron transport. A possible method for self-consistently modelling the distribution functions without requiring PIC simulations is discussed.

Investigation of excited states populations density of Hall thruster plasma in three dimensions by laser-induced fluorescence spectroscopy

Science.gov (United States)

Krivoruchko, D. D.; Skrylev, A. V.

2018-01-01

The article deals with investigation of the excited states populations distribution of a low-temperature xenon plasma in the thruster with closed electron drift at 300 W operating conditions were investigated by laser-induced fluorescence (LIF) over the 350-1100 nm range. Seven xenon ions (Xe II) transitions were analyzed, while for neutral atoms (Xe I) just three transitions were explored, since the majority of Xe I emission falls into the ultraviolet or infrared part of the spectrum and are difficult to measure. The necessary spontaneous emission probabilities (Einstein coefficients) were calculated. Measurements of the excited state distribution were made for points (volume of about 12 mm3) all over the plane perpendicular to thruster axis in four positions on it (5, 10, 50 and 100 mm). Measured LIF signal intensity have differences for each location of researched point (due to anisotropy of thruster plume), however the structure of states populations distribution persisted at plume and is violated at the thruster exit plane and cathode area. Measured distributions show that for describing plasma of Hall thruster one needs to use a multilevel kinetic model, classic model can be used just for far plume region or for specific electron transitions.

Detection of metal ions by atomic emission spectroscopy from liquid-electrode discharge plasma

International Nuclear Information System (INIS)

Wu Jian; Yu Jing; Li Jun; Wang Jianping; Ying Yibin

2007-01-01

In this paper, the discharge ignited in a capillary connecting two beakers filled with electrolyte solution is investigated. During the experiment, an external electrical voltage is applied through two platinum electrodes dipped in the beakers. A gas bubble forms inside the capillary when the applied voltage is higher than 1000 V. Since the beakers are tilted slightly, after generation, the bubble moves slowly to the uphill outlet of the capillary due to buoyancy. When the bubble reaches the end of the capillary, it cracks and a bright discharge is ignited. The emission spectra of the discharge plasma are related to the metal ions dissolved in the solution and thus can be used for metal ion detection. An application of the system to measurement of water hardness is shown

Development of high efficiency Versatile Arc Discharge Ion Source at CERN ISOLDE.

Science.gov (United States)

Penescu, L; Catherall, R; Lettry, J; Stora, T

2010-02-01

We report here recent developments of Forced Electron Beam Induced Arc Discharge (FEBIAD) ion sources at the ISOLDE radioactive ion beam facility, hosted at the European Organization for Nuclear Research (CERN). As a result of the propositions to improve the ionization efficiency, two FEBIAD prototypes have been produced and successfully tested in 2008. Off-line studies showed that the 1+ ionization efficiencies for noble gases are 5-20 times larger than with the standard ISOLDE FEBIAD ion sources and reach 60% for radon, which allowed the identification at ISOLDE of (229)Rn, an isotope that had never previously been observed in the laboratory. A factor of 3 increase is also expected for the ionization efficiency of the other elements. The experimental and theoretical methodology is presented. The theoretical model, which gives precise insights on the processes affecting the ionization, is used to design optimal sources (grouped under the name of VADIS--Versatile Arc Discharge Ion Source) for the different chemical classes of the produced isotopes, as already demonstrated for the noble gases.

On the role of ion heating in ICRF-heated discharges in Tore Supra

International Nuclear Information System (INIS)

Eriksson, L.G.; Hoang, G.; Bergeaud, V.

2000-09-01

The effect of bulk ion heating in Tore Supra has been investigated by studying discharges with varying concentrations of minority ions during ICRF hydrogen minority heating in Deuterium/ 4 He plasmas. As expected, the level of bulk ion heating is found to increase with the minority concentration. Higher levels of ion heating are shown to be accompanied by two significant effects: an improved energy confinement and a strong influence on the plasma rotation. (author)

High Accuracy Positioning using Jet Thrusters for Quadcopter

Directory of Open Access Journals (Sweden)

Pi ChenHuan

2018-01-01

Full Text Available A quadcopter is equipped with four additional jet thrusters on its horizontal plane and vertical to each other in order to improve the maneuverability and positioning accuracy of quadcopter. A dynamic model of the quadcopter with jet thrusters is derived and two controllers are implemented in simulation, one is a dual loop state feedback controller for pose control and another is an auxiliary jet thruster controller for accurate positioning. Step response simulations showed that the jet thruster can control the quadcopter with less overshoot compared to the conventional one. Over 10s loiter simulation with disturbance, the quadcopter with jet thruster decrease 85% of RMS error of horizontal disturbance compared to a conventional quadcopter with only a dual loop state feedback controller. The jet thruster controller shows the possibility for further accurate in the field of quadcopter positioning.

Electron and Negative Ion Densities in C2F6 and CHF3 Containing Inductively Coupled Discharges

International Nuclear Information System (INIS)

HEBNER, GREGORY A.; MILLER, PAUL A.

1999-01-01

Electron and negative ion densities have been measured in inductively coupled discharges containing C 2 F 6 and CHF 3 . Line integrated electron density was determined using a microwave interferometer, negative ion densities were inferred using laser photodetachment spectroscopy, and electron temperature was determined using a Langmuir probe. For the range of induction powers, pressures and bias power investigated, the electron density peaked at 9 x 10 12 cm -2 (line-integrated) or approximately 9 x 10 11 cm -3 . The negative ion density peaked at approximately 1.3 x 10 11 cm -3 . A maximum in the negative ion density as a function of induction coil power was observed. The maximum is attributed to a power dependent change in the density of one or more of the potential negative ion precursor species since the electron temperature did not depend strongly on power. The variation of photodetachment with laser wavelength indicated that the dominant negative ion was F - . Measurement of the decay of the negative ion density in the afterglow of a pulse modulated discharge was used to determine the ion-ion recombination rate for CF 4 , C 2 F 6 and CHF 3 discharges

Ion source using a hollow cathode discharge system and especially, particle accelerator comprising said source

International Nuclear Information System (INIS)

Mourier, Georges.

1975-01-01

An ion source provided with a hollow cathode discharge system is presented. The ion extraction system is designed in view of generating a beam directed towards a point of use located far from the point of ion production. Said source essentially comprises two cathodes facing each other, an anode at a continuous voltage with respect to the cathodes, a heated filament beyond the cathode on the path of the extracted beam, and a grid between said filament and cathode. The ion extraction is limited to a certain portion of the ions present inside the plasma, so as the discharge to continue to be sustained by itself. For that purpose pierced cathodes are used, with a transparency (the ratio of the hole area to the whole cathode area) not much higher than 50% [fr

Organic impurities harmful to arc discharge in multi-ampere duo-pigatron ion source

International Nuclear Information System (INIS)

Peng Zhudong; Wang Shouhu; Chen Miaosun; Huang Chongan; Ge Yishan; Fu Peng; E Maohuai

1989-01-01

Besides mechanism producing plasmain Duo-Pigatron Ion Source, the main purpose of this paper is to introduce the problem on arc discharge exhaustion caused by oil pollution. The problem was discovered during developing a 15 cm Duo-Pigatron Ion Source. The paper also describes its physical process

Hall effect thruster with an AlN chamber

International Nuclear Information System (INIS)

Barral, S.; Jayet, Y.; Mazouffre, S.; Veron, E.; Echegut, P.; Dudeck, M.

2005-01-01

The plasma discharge of a Hall-effect thruster (SPT) is strongly depending of the plasma-insulated wall interactions. These interactions are mainly related to the energy deposition, potential sheath effect and electron secondary emission rate (e.s.e.). In usual SPT, the annular channel is made of BN-SiO 2 . The SPT100-ML (laboratory model will be tested with an AlN chamber in the French test facility Pivoine in the laboratoire d'Aerothermique (Orleans-France). The different parameters such as discharge current, thrust, plasma oscillations and wall temperature will studied for several operating conditions. The results will be compared with a fluid model developed in IPPT (Warsaw-Poland) taking into account electron emission from the internal and external walls and using previous experimental measurements of e.s.e. for AlN from ONERA (Toulouse-France). The surface state of AlN will be analysed before and after experiments by an Environmental Scanning Electron Microscope and by a Strength Electron Microscope. (author)

Coaxial plasma thrusters for high specific impulse propulsion

Science.gov (United States)

Schoenberg, Kurt F.; Gerwin, Richard A.; Barnes, Cris W.; Henins, Ivars; Mayo, Robert; Moses, Ronald, Jr.; Scarberry, Richard; Wurden, Glen

1991-01-01

A fundamental basis for coaxial plasma thruster performance is presented and the steady-state, ideal MHD properties of a coaxial thruster using an annular magnetic nozzle are discussed. Formulas for power usage, thrust, mass flow rate, and specific impulse are acquired and employed to assess thruster performance. The performance estimates are compared with the observed properties of an unoptimized coaxial plasma gun. These comparisons support the hypothesis that ideal MHD has an important role in coaxial plasma thruster dynamics.

Laser injection of ultra-short electron bursts for the diagnosis of Hall thruster plasma

International Nuclear Information System (INIS)

Albarede, L; Gibert, T; Lazurenko, A; Bouchoule, A

2006-01-01

The present developments of Hall thrusters for satellite control and space mission technologies represent a new step towards their routine use in place of conventional thermal thrusters. In spite of their long R and D history, the complex physics of the E x B discharge at work in these structures has prevented, up to now, the availability of predictive simulations. The electron transport in the accelerating layers of these thrusters is one of the remaining challenges in this direction. From the experimental point of view, any diagnostics of electron transport and electric field in this critical layer would be welcome for comparison with code predictions. Appropriate diagnostics are difficult, due to the very aggressive local plasma conditions. This paper presents the first step in the development of a new tool for characterization of the plasma electric field in the very near exhaust thruster plume and comparison with simulation code predictions. The main idea is to use very short bursts of electrons, probing local electron dynamics in this critical plume area. Such bursts can be obtained through photoelectric emission induced by a UV pulsed laser beam on a convenient target. A specific study, devoted to the characterization of the electron burst emission, is presented in the first section of the paper; the implementation and testing of the injection of electrons in the critical layer of Hall thruster plasma is described in the second section. The design and testing of a fast and sensitive system for characterizing the transport of injected bursts will be the next step of this program. It requires a preliminary evaluation of electron trajectories which was achieved by using simulation code. Simulation data are presented in the last section of the paper, with the full diagnostic design to be tested in the near future, when runs will be available in the renewed PIVOINE facility. The same electron burst injection could also be a valuable input in the present

Magnetic Electron Filtering by Fluid Models for the PEGASES Thruster

Science.gov (United States)

Leray, Gary; Chabert, Pascal; Lichtenberg, Allan; Lieberman, Michael

2009-10-01

The PEGASES thruster produces thrust by creating positive and negative ions, which are then accelerated. To accelerate both type of ions, electrons need to be filtered, which is achieved by applying a static magnetic field strong enough to magnetize the electrons but not the ions. A 1D fluid model with three species (electrons, positive and negative ions) and an analytical model are proposed to understand this process for an oxygen plasma with p = 10 mTorr and B0 = 300 G [1]. The resulting ion-ion plasma formation in the transverse direction (perpendicular to the magnetic field) is demonstrated. It is shown that an additional electron/positive ion loss term is required. The solutions are evaluated for two main parameters: the ionizing fraction at the plasma center (x = 0), ne0/ng, and the electronegativity ratio at the center, α0=nn0/ne0. The effect of geometry and magnetic field amplitude are also discussed. [4pt] [1] Leray G, Chabert P, Lichtenberg A J and Lieberman M A, J. Phys. D: Appl. Phys., Plasma Modelling Cluster issue, to appear (2009)

Performance and flow characteristics of MHD seawater thruster

Energy Technology Data Exchange (ETDEWEB)

Doss, E.D.

1990-01-01

The main goal of the research is to investigate the effects of strong magnetic fields on the electrical and flow fields inside MHD thrusters. The results of this study is important in the assessment of the feasibility of MHD seawater propulsion for the Navy. To accomplish this goal a three-dimensional fluid flow computer model has been developed and applied to study the concept of MHD seawater propulsion. The effects of strong magnetic fields on the current and electric fields inside the MHD thruster and their interaction with the flow fields, particularly those in the boundary layers, have been investigated. The results of the three-dimensional computations indicate that the velocity profiles are flatter over the sidewalls of the thruster walls in comparison to the velocity profiles over the electrode walls. These nonuniformities in the flow fields give rise to nonuniform distribution of the skin friction along the walls of the thrusters, where higher values are predicted over the sidewalls relative to those over the electrode walls. Also, a parametric study has been performed using the three-dimensional MHD flow model to analyze the performance of continuous electrode seawater thrusters under different operating parameters. The effects of these parameters on the fluid flow characteristics, and on the thruster efficiency have been investigated. Those parameters include the magnetic field (10--20 T), thruster diameter, surface roughness, flow velocity, and the electric load factor. The results show also that the thruster performance improves with the strength of the magnetic field and thruster diameter, and the efficiency decreases with the flow velocity and surface roughness.

Plasma-Sheath Instability in Hall Thrusters Due to Periodic Modulation of the Energy of Secondary Electrons in Cyclotron Motion

International Nuclear Information System (INIS)

Sydorenko, D.; Smolyakov, A.; Kaganovich, I.; Raitses, Y.

2008-01-01

Particle-in-cell simulation of Hall thruster plasmas reveals a plasma-sheath instability manifesting itself as a rearrangement of the plasma sheath near the thruster channel walls accompanied by a sudden change of many discharge parameters. The instability develops when the sheath current as a function of the sheath voltage is in the negative conductivity regime. The major part of the sheath current is produced by beams of secondary electrons counter-streaming between the walls. The negative conductivity is the result of nonlinear dependence of beam-induced secondary electron emission on the plasma potential. The intensity of such emission is defined by the beam energy. The energy of the beam in crossed axial electric and radial magnetic fields is a quasi-periodical function of the phase of cyclotron rotation, which depends on the radial profile of the potential and the thruster channel width. There is a discrete set of stability intervals determined by the final phase of the cyclotron rotation of secondary electrons. As a result, a small variation of the thruster channel width may result in abrupt changes of plasma parameters if the plasma state jumps from one stability interval to another

Study and Developement of Compact Permanent Magnet Hall Thrusters for Future Brazillian Space Missions

Science.gov (United States)

Ferreira, Jose Leonardo; Martins, Alexandre; Cerda, Rodrigo

2016-07-01

. The main difficulty to reach these minor bodies is related to their specific orbits with high eccentricity and inclination. A good example is the case for sample return missions to NEOs-Near Earth Objects. They are small bodies consisting of primitive left over building blocks of the Solar System formation processes. These missions can be accomplished by using low thrust trajectories with spacecrafts propelled by plasma thrusters with total thrust below 0.5 N, and a specific impulse around2500 s. In this work, we will show the brazilian contribution to the development of a compact electrical propulsion engine named PHALL III, designed with DCFH and foreseen to be used in future cubesats microsatellites but with possible applications in geostationary attitude control systems and on low thrust trajectory missions to the Near Earth Asteroids region. We will show a particular new permanent magnet field designed for PHALL III . Computer based simulation codes such as VSIM are also used on the design of this new proposed cuped magnet field Hall Thruster. Based on the first results wee believed PHALL III will also allow a good spacecraft performance of long duration space missions for small size spacecrafts with limited low electric source power consumption. The PHALL III plasma source characterization is presented together with the ejected plasma plume ion current intensity, ion energy and plasma flow velocity parameters measured by an integrated Plasma Diagnostic Bench (BID). Based on plasma source and plume ejected parameters a merit figure of PHALL III is constructed and compared to computer calculated low thrust transfer requirements. From these results it is goig to be possible to analyse the potential use of PHALL III on future brazillian space missions , its working parameters are compared with parameters of existing space tested plasma thrusters already used on moon , deep space missions and also on satellite geostationary positioning using low thrust orbit

Ion extraction capabilities of two-grid accelerator systems

International Nuclear Information System (INIS)

Rovang, D.C.; Wilbur, P.J.

1984-02-01

An experimental investigation into the ion extraction capabilities of two-grid accelerator systems common to electrostatic ion thrusters is described. This work resulted in a large body of experimental data which facilitates the selection of the accelerator system geometries and operating parameters necessary to maximize the extracted ion current. Results suggest that the impingement-limited perveance is not dramatically affected by reductions in screen hole diameter to 0.5 mm. Impingement-limited performance is shown to depend most strongly on grid separation distance, accelerator hole diameter ratio, the discharge-to-total accelerating voltage ratio, and the net-to-total accelerating voltage ratio. Results obtained at small grid separation ratios suggest a new grid operating condition where high beam current per hole levels are achieved at a specified net accelerating voltage. It is shown that this operating condition is realized at an optimum ratio of net-to-total accelerating voltage ratio which is typically quite high. The apparatus developed for this study is also shown to be well suited measuring the electron backstreaming and electrical breakdown characteristics of two-grid accelerator systems

Ignition method of corona discharge with modulation of the field in ion source of ion mobility spectrometer

International Nuclear Information System (INIS)

Gromov, Evgeniy

2011-01-01

The new method for the ignition of the corona discharge has been developed, which improves the stability of the ion mobility spectrometer and the resolution of the instrument. The system of forming a corona discharge without additional electrodes, which are used in a number of known structures for the pre-ionization, has been developed. This simplifies the design of the proposed source and an electronic control circuit. IMS technology is widely used in different civil and military fields for vapor-phase detection of explosive, narcotics, chemical warfare agents, biology molecules and so on. There are set of methods whose are used for the ionization of molecules under analysis. They are the following: radioactive ionization, ultraviolet photoionization, laser ionization, electric field ionization, corona spray ionization, electro spray ionization, roentgen ionization, and surface ionization. All these methods has their own advantages and disadvantages. A comparing of ion mobility spectra of non-polar hydrocarbons for photoionization, corona discharge ionization and 63 Ni ionization, had carried in. In our work we have investigated four types of IMS spectrometers whose use different sources for molecules under analysis ionization. They use radioactive ionization, ultraviolet photoionization, laser ionization, and roentgen ionization. The traditional explosives had investigated in experiments. In electricity, a corona discharge is an electrical discharge brought on by the ionization of a fluid surrounding a conductor, which occurs when the potential gradient (the strength of the electric field) exceeds a certain value, but conditions are insufficient to cause complete electrical breakdown or arcing.

Pressure History Measurement in a Microwave Beaming Thruster

International Nuclear Information System (INIS)

Oda, Yasuhisa; Ushio, Masato; Komurasaki, Kimiya; Takahashi, Koji; Kasugai, Atsushi; Sakamoto, Keishi

2006-01-01

In a microwave beaming thruster with a 1-dimensional nozzle, plasma and shock wave propagates in the nozzle absorbing microwave power. In this study, pressure histories in the thruster are measured using pressure gauges. Measured pressure history at the thruster wall shows constant pressure during plasma propagation in the nozzle. The result of measurement of the propagating velocities of shock wave and plasma shows that both propagate in the same velocity. These result shows that thrust producing model of analogy of pulse detonation engine is successful for the 1D thruster

Diagnostics of a negative hydrogen ion source based on a planar-coil inductively-driven discharge

International Nuclear Information System (INIS)

Lishev, Stiliyan St.; Shivarova, Antonia P.; Iordanov, Dimitar I.; Todorov, Dimitar T.; Demerdzhiev, Angel P.

2013-01-01

Recent activity on a volume-production negative-hydrogen-ion source with the design of a matrix of small-radius discharges inductively driven by a planar coil is reported. A single discharge of the matrix has been studied by employing probe-and laser-photodetachment diagnostics. The measurements have been carried out in the first - small-size - chamber of a two-chamber discharge vessel, with rf power introduced by a planar coil positioned on its front wall. The measured profiles of the negative ion density show strong axial dependence, with two maxima located at the positions corresponding to the maxima of the dc potential, where the maxima of the electron density are also located. The first maximum is in the vicinity of the rf power input and the second one is close to the interface between the two chambers of the discharge vessel.

A cavity ring-down spectroscopy sensor for real-time Hall thruster erosion measurements

International Nuclear Information System (INIS)

Lee, B. C.; Huang, W.; Tao, L.; Yamamoto, N.; Yalin, A. P.; Gallimore, A. D.

2014-01-01

A continuous-wave cavity ring-down spectroscopy sensor for real-time measurements of sputtered boron from Hall thrusters has been developed. The sensor uses a continuous-wave frequency-quadrupled diode laser at 250 nm to probe ground state atomic boron sputtered from the boron nitride insulating channel. Validation results from a controlled setup using an ion beam and target showed good agreement with a simple finite-element model. Application of the sensor for measurements of two Hall thrusters, the H6 and SPT-70, is described. The H6 was tested at power levels ranging from 1.5 to 10 kW. Peak boron densities of 10 ± 2 × 10 14 m −3 were measured in the thruster plume, and the estimated eroded channel volume agreed within a factor of 2 of profilometry. The SPT-70 was tested at 600 and 660 W, yielding peak boron densities of 7.2 ± 1.1 × 10 14 m −3 , and the estimated erosion rate agreed within ∼20% of profilometry. Technical challenges associated with operating a high-finesse cavity in the presence of energetic plasma are also discussed

A Numerical Study on Hydrodynamic Interactions between Dynamic Positioning Thrusters

Energy Technology Data Exchange (ETDEWEB)

Jin, Doo Hwa; Lee, Sang Wook [University of Ulsan, Ulsan (Korea, Republic of)

2017-06-15

In this study, we conducted computational fluid dynamics (CFD) simulations for the unsteady hydrodynamic interaction of multiple thrusters by solving Reynolds averaged Navier-Stokes equations. A commercial CFD software, STAR-CCM+ was used for all simulations by employing a ducted thruster model with combination of a propeller and No. 19a duct. A sliding mesh technique was used to treat dynamic motion of propeller rotation and non-conformal hexahedral grid system was considered. Four different combinations in tilting and azimuth angles of the thrusters were considered to investigate the effects on the propulsion performance. We could find that thruster-hull and thruster-thruster interactions has significant effect on propulsion performance and further study will be required for the optimal configurations with the best tilting and relative azimuth angle between thrusters.

Arcjet space thrusters

Science.gov (United States)

Keefer, Dennis; Rhodes, Robert

1993-05-01

Electrically powered arc jets which produce thrust at high specific impulse could provide a substantial cost reduction for orbital transfer and station keeping missions. There is currently a limited understanding of the complex, nonlinear interactions in the plasma propellant which has hindered the development of high efficiency arc jet thrusters by making it difficult to predict the effect of design changes and to interpret experimental results. A computational model developed at the University of Tennessee Space Institute (UTSI) to study laser powered thrusters and radio frequency gas heaters has been adapted to provide a tool to help understand the physical processes in arc jet thrusters. The approach is to include in the model those physical and chemical processes which appear to be important, and then to evaluate our judgement by the comparison of numerical simulations with experimental data. The results of this study have been presented at four technical conferences. The details of the work accomplished in this project are covered in the individual papers included in the appendix of this report. We present a brief description of the model covering its most important features followed by a summary of the effort.

Pilot scale ion exchange column study for reducing radioactivity discharges to environment

International Nuclear Information System (INIS)

Kore, S.G.; Yadav, V.K.; Sonar, N.L.; Valsala, T.P.; Narayan, J.; Sharma, S.P.; Chattopadhyay, S.; Dani, U.; Vishwaraj, I.

2013-01-01

Low level liquid waste (LLW) is generated during operation of Tarapur Atomic Power Station (TAPS). Chemical co-precipitation is the treatment method used for decontamination of this waste with respect to radionuclide prior to discharge to environment. Further polishing of effluent from the treated LLW was planned using ion exchange column to reduce the discharges to the environment In view of this ion exchange column study was carried out in the laboratory using in-house prepared cobalt ferrocyanide (COFC) based composite resin. Based on the encouraging results obtained in the lab studies, pilot scale study was carried out in the plant. Decontamination factor (DF) of 14-15 was obtained with respect to Cs isotopes and overall DF of 2-5 was obtained with respect to gross beta activity. (author)

The discharge behavior of lithium-ion batteries using the Dual-Potential Multi-Scale Multi-Dimensional (MSMD) Battery Model

DEFF Research Database (Denmark)

Saeed Madani, Seyed; Swierczynski, Maciej Jozef; Kær, Søren Knudsen

2017-01-01

This paper gives insight into the discharge behavior of lithium-ion batteries based on the investigations, which have been done by the researchers [1– 19]. In this article, the battery's discharge behaviour at various discharge rates is studied and surface monitor, discharge curve, volume monitor...... to analysis the discharge behaviour of lithium-ion batteries. The results show that surface monitor plot of discharge curve at 1 C has a decreasing trend and volume monitor plot of maximum temperature in the domain has slightly increasing pattern over the simulation time. For the curves of discharge...... plot of maximum temperature in the domain and maximum temperature in the area are illustrated. Additionally, an external and internal short-circuit treatment for three cases have been studied. The Dual-Potential Multi-Scale Multi-Dimensional (MSMD) Battery Model (BM) was used by ANSYS FLUENT software...

Trade Study of Multiple Thruster Options for the Mars Airplane Concept

Science.gov (United States)

Kuhl, Christopher A.; Gayle, Steven W.; Hunter, Craig A.; Kenney, Patrick S.; Scola, Salvatore; Paddock, David A.; Wright, Henry S.; Gasbarre, Joseph F.

2009-01-01

A trade study was performed at NASA Langley Research Center under the Planetary Airplane Risk Reduction (PARR) project (2004-2005) to examine the option of using multiple, smaller thrusters in place of a single large thruster on the Mars airplane concept with the goal to reduce overall cost, schedule, and technical risk. The 5-lbf (22N) thruster is a common reaction control thruster on many satellites. Thousands of these types of thrusters have been built and flown on numerous programs, including MILSTAR and Intelsat VI. This study has examined the use of three 22N thrusters for the Mars airplane propulsion system and compared the results to those of the baseline single thruster system.

Generation of ion-acoustic waves in an inductively coupled, low-pressure discharge lamp

International Nuclear Information System (INIS)

Camparo, J. C.; Klimcak, C. M.

2006-01-01

For a number of years it has been known that the alkali rf-discharge lamps used in atomic clocks can exhibit large amplitude intensity oscillations. These oscillations arise from ion-acoustic plasma waves and have typically been associated with erratic clock behavior. Though large amplitude ion-acoustic plasma waves are clearly deleterious for atomic clock operation, it does not follow that small amplitude oscillations have no utility. Here, we demonstrate two easily implemented methods for generating small amplitude ion-acoustic plasma waves in alkali rf-discharge lamps. Furthermore, we demonstrate that the frequency of these waves is proportional to the square root of the rf power driving the lamp and therefore that their examination can provide an easily accessible parameter for monitoring and controlling the lamp's plasma conditions. This has important consequences for precise timekeeping, since the atomic ground-state hyperfine transition, which is the heart of the atomic clock signal, can be significantly perturbed by changes in the lamp's output via the ac-Stark shift

Ion energy and angular distributions in inductively coupled Argon RF discharges

International Nuclear Information System (INIS)

Woodworth, J.R.; Riley, M.E.; Meister, D.C.

1996-03-01

We report measurements of the energies and angular distributions of positive ions in an inductively coupled argon plasma in a GEC reference cell. Use of two separate ion detectors allowed measurement of ion energies and fluxes as a function of position as well as ion angular distributions on the discharge centerline. The inductive drive on our system produced high plasma densities (up to 10 12 /cm 3 electron densities) and relatively stable plasma potentials. As a result, ion energy distributions typically consisted of a single feature well separated from zero energy. Mean ion energy was independent of rf power and varied inversely with pressure, decreasing from 29 eV to 12 eV as pressure increased form 2.4 m Torr to 50 mTorr. Half-widths of the ion angular distributions in these experiments varied from 5 degrees to 12.5 degrees, or equivalently, transverse temperatures varied form 0.2 to 0.5 eV with the distributions broadening as either pressure or RF power were increased

Mechanical design of SERT 2 thruster system

Science.gov (United States)

Zavesky, R. J.; Hurst, E. B.

1972-01-01

The mechanical design of the mercury bombardment thruster that was tested on SERT is described. The report shows how the structural, thermal, electrical, material compatibility, and neutral mercury coating considerations affected the design and integration of the subsystems and components. The SERT 2 spacecraft with two thrusters was launched on February 3, 1970. One thruster operated for 3782 hours and the other for 2011 hours. A high voltage short resulting from buildup of loose eroded material was believed to be the cause of failure.

Two-dimensional magnetic field evolution measurements and plasma flow speed estimates from the coaxial thruster experiment

International Nuclear Information System (INIS)

Black, D.C.; Mayo, R.M.; Gerwin, R.A.; Schoenberg, K.F.; Scheuer, J.T.; Hoyt, R.P.; Henins, I.

1994-01-01

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 v z ∼10 7 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 distortion by the highly conductive plasma is occurring, and to provide insight into the details of discharge evolution. Data from a magnetic fluctuation probe array have been used to investigate the details of applied magnetic field deformation through the reconstruction of time-dependent flux profiles. Experimentally observed magnetic field line distortion has been compared to that predicted by a simple one-dimensional (1-D) model of the discharge channel. Such a comparison is utilized to estimate the axial plasma velocity in the thruster. Velocities determined in this manner are in approximate agreement with the predicted self-field magnetosonic speed and those measured by a time-of-flight spectrometer

Control of the electric-field profile in the Hall thruster

International Nuclear Information System (INIS)

Fruchtman, A.; Fisch, N.J.; Raitses, Y.

2001-01-01

Control of the electric-field profile in the Hall thruster through the positioning of an additional electrode along the channel is shown theoretically to enhance the efficiency. The reduction of the potential drop near the anode by use of the additional electrode increases the plasma density there, through the increase of the electron and ion transit times, causing the ionization in the vicinity of the anode to increase. The resulting separation of the ionization and acceleration regions increases the propellant and energy utilizations. An abrupt sonic transition is forced to occur at the axial location of the additional electrode, accompanied by the generation of a large (theoretically infinite) electric field. This ability to generate a large electric field at a specific location along the channel, in addition to the ability to specify the electric potential there, allows us further control of the electric-field profile in the thruster. In particular, when the electron temperature is high, a large abrupt voltage drop is induced at the vicinity of the additional electrode, a voltage drop that can comprise a significant part of the applied voltage

Studies of Non-Conventional Configuration Closed Electron Drift Thrusters

International Nuclear Information System (INIS)

Y. Raitses; D. Staack; A. Smirnov; A.A. Litvak; L.A. Dorf; T. Graves; N.J. Fisch

2001-01-01

In this paper, we review recent results obtained for segmented electrode and cylindrical Hall thrusters. A low sputtering graphite segmented electrode, placed at the exit of the annular thruster, is shown to affect the plasma potential distribution in the ceramic channel. This effect appears to be correlated with an observed plume reduction compared to a conventional, nonsegmented thruster. In preliminary experiments a 3-cm thruster was operated in the 50-200 W power range. Two operating regimes, stable and oscillating, were observed and investigated

System analysis and test-bed for an atmosphere-breathing electric propulsion system using an inductive plasma thruster

Science.gov (United States)

Romano, F.; Massuti-Ballester, B.; Binder, T.; Herdrich, G.; Fasoulas, S.; Schönherr, T.

2018-06-01

Challenging space mission scenarios include those in low altitude orbits, where the atmosphere creates significant drag to the S/C and forces their orbit to an early decay. For drag compensation, propulsion systems are needed, requiring propellant to be carried on-board. An atmosphere-breathing electric propulsion system (ABEP) ingests the residual atmosphere particles through an intake and uses them as propellant for an electric thruster. Theoretically applicable to any planet with atmosphere, the system might allow to orbit for unlimited time without carrying propellant. A new range of altitudes for continuous operation would become accessible, enabling new scientific missions while reducing costs. Preliminary studies have shown that the collectible propellant flow for an ion thruster (in LEO) might not be enough, and that electrode erosion due to aggressive gases, such as atomic oxygen, will limit the thruster lifetime. In this paper an inductive plasma thruster (IPT) is considered for the ABEP system. The starting point is a small scale inductively heated plasma generator IPG6-S. These devices are electrodeless and have already shown high electric-to-thermal coupling efficiencies using O2 and CO2 . The system analysis is integrated with IPG6-S tests to assess mean mass-specific energies of the plasma plume and estimate exhaust velocities.

Evaluation of ion collection area in Faraday probes.

Science.gov (United States)

Brown, Daniel L; Gallimore, Alec D

2010-06-01

A Faraday probe with three concentric rings was designed and fabricated to assess the effect of gap width and collector diameter in a systematic study of the diagnostic ion collection area. The nested Faraday probe consisted of two concentric collector rings and an outer guard ring, which enabled simultaneous current density measurements on the inner and outer collectors. Two versions of the outer collector were fabricated to create gaps of 0.5 and 1.5 mm between the rings. Distribution of current density in the plume of a low-power Hall thruster ion source was measured in azimuthal sweeps at constant radius from 8 to 20 thruster diameters downstream of the exit plane with variation in facility background pressure. A new analytical technique is proposed to account for ions collected in the gap between the Faraday probe collector and guard ring. This method is shown to exhibit excellent agreement between all nested Faraday probe configurations, and to reduce the magnitude of integrated ion beam current to levels consistent with Hall thruster performance analyses. The technique is further studied by varying the guard ring bias potential with a fixed collector bias potential, thereby controlling ion collection in the gap. Results are in agreement with predictions based on the proposed analytical technique. The method is applied to a past study comparing the measured ion current density profiles of two Faraday probe designs. These findings provide new insight into the nature of ion collection in Faraday probe diagnostics, and lead to improved accuracy with a significant reduction in measurement uncertainty.

Evaluation of ion collection area in Faraday probes

International Nuclear Information System (INIS)

Brown, Daniel L.; Gallimore, Alec D.

2010-01-01

A Faraday probe with three concentric rings was designed and fabricated to assess the effect of gap width and collector diameter in a systematic study of the diagnostic ion collection area. The nested Faraday probe consisted of two concentric collector rings and an outer guard ring, which enabled simultaneous current density measurements on the inner and outer collectors. Two versions of the outer collector were fabricated to create gaps of 0.5 and 1.5 mm between the rings. Distribution of current density in the plume of a low-power Hall thruster ion source was measured in azimuthal sweeps at constant radius from 8 to 20 thruster diameters downstream of the exit plane with variation in facility background pressure. A new analytical technique is proposed to account for ions collected in the gap between the Faraday probe collector and guard ring. This method is shown to exhibit excellent agreement between all nested Faraday probe configurations, and to reduce the magnitude of integrated ion beam current to levels consistent with Hall thruster performance analyses. The technique is further studied by varying the guard ring bias potential with a fixed collector bias potential, thereby controlling ion collection in the gap. Results are in agreement with predictions based on the proposed analytical technique. The method is applied to a past study comparing the measured ion current density profiles of two Faraday probe designs. These findings provide new insight into the nature of ion collection in Faraday probe diagnostics, and lead to improved accuracy with a significant reduction in measurement uncertainty.

Time resolved energy spectrum of the axial ion beam generated in plasma focus discharges

International Nuclear Information System (INIS)

Bostick, W.H.; Kilic, H.; Nardi, V.; Powell, C.W.

1993-01-01

The energy spectrum of the deuteron beam along the electrode axis (0 (degree) ) in a plasma focus discharge has been determined with a time of flight (TOF) method and with a differential filter method in the ion energy interval E = 0.3-9 MeV. The ion TOF method is applied to single-ion pulse events with an ion emission time t(E) that is only weakly dependent on the ion energy E for E > 0.3 MeV. The correlation of the ion beam intensity with the filling pressure, the neutron yield and the hard X-ray intensity is also reported. (author). 11 refs, 10 figs

Engineering Risk Assessment of Space Thruster Challenge Problem

Science.gov (United States)

Mathias, Donovan L.; Mattenberger, Christopher J.; Go, Susie

2014-01-01

The Engineering Risk Assessment (ERA) team at NASA Ames Research Center utilizes dynamic models with linked physics-of-failure analyses to produce quantitative risk assessments of space exploration missions. This paper applies the ERA approach to the baseline and extended versions of the PSAM Space Thruster Challenge Problem, which investigates mission risk for a deep space ion propulsion system with time-varying thruster requirements and operations schedules. The dynamic mission is modeled using a combination of discrete and continuous-time reliability elements within the commercially available GoldSim software. Loss-of-mission (LOM) probability results are generated via Monte Carlo sampling performed by the integrated model. Model convergence studies are presented to illustrate the sensitivity of integrated LOM results to the number of Monte Carlo trials. A deterministic risk model was also built for the three baseline and extended missions using the Ames Reliability Tool (ART), and results are compared to the simulation results to evaluate the relative importance of mission dynamics. The ART model did a reasonable job of matching the simulation models for the baseline case, while a hybrid approach using offline dynamic models was required for the extended missions. This study highlighted that state-of-the-art techniques can adequately adapt to a range of dynamic problems.

Ion energy distributions in bipolar pulsed-dc discharges of methane measured at the biased cathode

Energy Technology Data Exchange (ETDEWEB)

Corbella, C; Rubio-Roy, M; Bertran, E; Portal, S; Pascual, E; Polo, M C; Andujar, J L, E-mail: corbella@ub.edu [FEMAN Group, IN2UB, Departament de Fisica Aplicada i Optica, Universitat de Barcelona, c/ MartI i Franques 1, 08028 Barcelona (Spain)

2011-02-15

The ion fluxes and ion energy distributions (IED) corresponding to discharges in methane (CH{sub 4}) were measured in time-averaged mode with a compact retarding field energy analyser (RFEA). The RFEA was placed on a biased electrode at room temperature, which was powered by either radiofrequency (13.56 MHz) or asymmetric bipolar pulsed-dc (250 kHz) signals. The shape of the resulting IED showed the relevant populations of ions bombarding the cathode at discharge parameters typical in the material processing technology: working pressures ranging from 1 to 10 Pa and cathode bias voltages between 100 and 200 V. High-energy peaks in the IED were detected at low pressures, whereas low-energy populations became progressively dominant at higher pressures. This effect is attributed to the transition from collisionless to collisional regimes of the cathode sheath as the pressure increases. On the other hand, pulsed-dc plasmas showed broader IED than RF discharges. This fact is connected to the different working frequencies and the intense peak voltages (up to 450 V) driven by the pulsed power supply. This work improves our understanding in plasma processes at the cathode level, which are of crucial importance for the growth and processing of materials requiring controlled ion bombardment. Examples of industrial applications with these requirements are plasma cleaning, ion etching processes during fabrication of microelectronic devices and plasma-enhanced chemical vapour deposition of hard coatings (diamond-like carbon, carbides and nitrides).

Pulsed Electrogasdynamic Thruster for Attitude Control and Orbit Maneuver, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — A new pulsed electric thruster, named "pulsed electrogasdynamic thruster," for attitude control and orbit maneuver is proposed. In this thruster, propellant gas is...

Magnetically Filtered Faraday Probe for Measuring the Ion Current Density Profile of a Hall Thruster

National Research Council Canada - National Science Library

Rovey, Joshua L; Walker, Mitchell L. R; Gallimore, Alec D; Peterson, Peter Y

2006-01-01

.../s. The probes are evaluated on a xenon propellant Hall thruster in the University of Michigan Large Vacuum Test Facility at operating pressures within the range of 4.4 x 10(-4) Pa Xe (3.3 x 10(-6) Torr Xe) to 1.1 10(-3) Pa Xe (8.4 x 10(-6) Torr Xe...

Development and characterization of high-efficiency, high-specific impulse xenon Hall thrusters

Science.gov (United States)

Hofer, Richard Robert

decrease of efficiency due to multiply-charged ions was minor. Efficiency was largely determined by the current utilization, which suggested maximum Hall thruster efficiency has yet to be reached. The electron Hall parameter was approximately constant with voltage, decreasing from an average of 210 at 300 V to an average of 160 between 400--900 V, which confirmed efficient operation can be realized only over a limited range of Hall parameters.

Atmospheric-pressure solution-cathode glow discharge: A versatile ion source for atomic and molecular mass spectrometry

Energy Technology Data Exchange (ETDEWEB)

Schwartz, Andrew J. [Department of Chemistry, Indiana University, Bloomington, IN, 47405 (United States); Williams, Kelsey L. [Department of Chemistry and Biochemistry, Kent State University, Kent, OH, 44242 (United States); Hieftje, Gary M. [Department of Chemistry, Indiana University, Bloomington, IN, 47405 (United States); Shelley, Jacob T., E-mail: shellj@rpi.edu [Department of Chemistry and Biochemistry, Kent State University, Kent, OH, 44242 (United States); Department of Chemistry and Chemical Biology, Rensselaer Polytechnic Institute, Troy, NY, 12180 (United States)

2017-01-15

An atmospheric-pressure solution-cathode glow discharge (SCGD) has been evaluated as an ion source for atomic, molecular, and ambient desorption/ionization mass spectrometry. The SCGD consists of a direct-current plasma, supported in the ambient air in the absence of gas flows, and sustained upon the surface of a flowing liquid cathode. Analytes introduced in the flowing liquid, as an ambient gas, or as a solid held near the plasma are vaporized and ionized by interactions within or near the discharge. Introduction of acidic solutions containing metal salts produced bare elemental ions as well as H{sub 2}O, OH{sup −} and NO{sub 3}{sup −} adducts. Detection limits for these elemental species ranged from 0.1 to 4 ppb, working curves spanned more than 4 orders of linear dynamic range, and precision varied between 5 and 16% relative standard deviation. Small organic molecules were also efficiently ionized from solution, and both the intact molecular ion and fragments were observed in the resulting SCGD mass spectra. Fragmentation of molecular species was found to be tunable; high discharge currents led to harder ionization, while low discharge currents produced stronger molecular-ion signals. Ambient gases and solids, desorbed by the plasma from a glass probe, were also readily ionized by the SCGD. Indeed, strong analyte signals were obtained from solid samples placed at least 2 cm from the plasma. These findings indicate that the SCGD might be useful also for ambient desorption/ionization mass spectrometry. Combined with earlier results that showed the SCGD is useful for ionization of labile biomolecules, the results here indicate that the SCGD is a highly versatile ion source capable of providing both elemental and molecular mass-spectral information. - Highlights: • Solution-cathode glow discharge used as an ionization source for mass spectrometry. • SCGD-MS can provide atomic as well as intact molecular mass spectra. • Atomic limits of detection range

Design and model experiments on thruster assisted mooring system; Futaishiki kaiyo kozobutsu no thruster ni yoru choshuki doyo seigyo

Energy Technology Data Exchange (ETDEWEB)

Nakamura, M; Koterayama, W [Kyushu Univ., Fukuoka (Japan). Research Inst. for Applied Mechanics; Kajiwara, H [Kyushu Institute of Technology, Kitakyushu (Japan). Faculty of Computer Science and System Engineering; Hyakudome, T [Kyushu University, Fukuoka (Japan)

1997-12-31

Described herein are dynamics and model experiments of the system in which positioning of a floating marine structure by mooring is combined with thruster-controlled positioning. Coefficients of dynamic forces acting on a floating structure model are determined experimentally and by the three-dimensional singularity distribution method, and the controller is designed by the PID, LQI and H{infinity} control theories. A model having a scale ratio of 1/100 was used for the experiments, where 2 thrusters were arranged in a diagonal line, one on the X-axis. It is found that the LQI and H{infinity} controllers of the thruster can control long-cycle rolling of the floating structure. They allow thruster control which is insensitive to wave cycle motion, and efficiently reduce positioning energy. The H{infinity} control regulates frequency characteristics of a closed loop more finely than the LQI control, and exhibits better controllability. 25 refs., 25 figs.

Design and model experiments on thruster assisted mooring system; Futaishiki kaiyo kozobutsu no thruster ni yoru choshuki doyo seigyo

Energy Technology Data Exchange (ETDEWEB)

Nakamura, M.; Koterayama, W. [Kyushu Univ., Fukuoka (Japan). Research Inst. for Applied Mechanics; Kajiwara, H. [Kyushu Institute of Technology, Kitakyushu (Japan). Faculty of Computer Science and System Engineering; Hyakudome, T. [Kyushu University, Fukuoka (Japan)

1996-12-31

Described herein are dynamics and model experiments of the system in which positioning of a floating marine structure by mooring is combined with thruster-controlled positioning. Coefficients of dynamic forces acting on a floating structure model are determined experimentally and by the three-dimensional singularity distribution method, and the controller is designed by the PID, LQI and H{infinity} control theories. A model having a scale ratio of 1/100 was used for the experiments, where 2 thrusters were arranged in a diagonal line, one on the X-axis. It is found that the LQI and H{infinity} controllers of the thruster can control long-cycle rolling of the floating structure. They allow thruster control which is insensitive to wave cycle motion, and efficiently reduce positioning energy. The H{infinity} control regulates frequency characteristics of a closed loop more finely than the LQI control, and exhibits better controllability. 25 refs., 25 figs.

Environmental Testing of the NEXT PM1 Ion Engine

Science.gov (United States)

Synder, John S.; Anderson, John R.; VanNoord, Jonathan L.; Soulas, George C.

2008-01-01

The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The Prototype Model engine PM1 was subjected to qualification-level environmental testing to demonstrate compatibility with environments representative of anticipated mission requirements. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 Grms for 2 minutes in each of three axes. Thermal-vacuum testing included a deep cold soak of the engine to temperatures of -168 C and thermal cycling from -120 to 203 C. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. Thruster performance was nominal throughout the test program, with minor variations in some engine operating parameters likely caused by facility effects. In general, the NEXT PM1 engine and the breadboard gimbal were found to be well-designed against environmental requirements based on the results reported herein. After resolution of the findings from this test program the hardware environmental qualification program can proceed with confidence.

Generation of ion-acoustic and magnetoacoustic waves in an RF helicon discharge

International Nuclear Information System (INIS)

Belov, A. S.; Markov, G. A.

2006-01-01

A study is made of the generation of ion-acoustic and magnetoacoustic waves in a discharge excited in an external magnetic field by an electromagnetic wave in the whistler frequency range (ω LH He , where ω LH = √(ω He ω Hi ) and ω He and ω Hi are the electron and ion gyrofrequencies, respectively). The excitation of acoustic waves is attributed to the decay of a high-frequency hybrid mode forming a plasma waveguide into low-frequency acoustic waves and new high-frequency waves that satisfy both the decay conditions and the waveguide dispersion relations. The excitation of acoustic waves is resonant in character because the conditions for the generation of waveguide modes and for the occurrence of the corresponding nonlinear wave processes should be satisfied simultaneously. An unexpected effect is the generation of magnetoacoustic waves by whistlers. A diagnostic technique is proposed that allows one to determine the thermal electron velocity by analyzing decay conditions and dispersion relations for waves in the discharge channel

Magnetic discharge accelerating diode for the gas-filled pulsed neutron generators based on inertial confinement of ions

International Nuclear Information System (INIS)

Kozlovskij, K I; Shikanov, A E; Vovchenko, E D; Shatokhin, V L; Isaev, A A; Martynenko, A S

2016-01-01

The paper deals with magnetic discharge diode module with inertial electrostatic ions confinement for the gas-filled pulsed neutron generators. The basis of the design is geometry with the central hollow cathode surrounded by the outer cylindrical anode and electrodes made of permanent magnets. The induction magnitude about 0.1-0.4 T in the central region of the discharge volume ensures the confinement of electrons in the space of hollow (virtual) cathode and leads to space charge compensation of accelerated ions in the centre. The research results of different excitation modes in pulsed high-voltage discharge are presented. The stable form of the volume discharge preserveing the shape and amplitude of the pulse current in the pressure range of 10 -3 -10 -1 Torr and at the accelerating voltage up to 200 kV was observed. (paper)

E × B electron drift instability in Hall thrusters: Particle-in-cell simulations vs. theory

Science.gov (United States)

Boeuf, J. P.; Garrigues, L.

2018-06-01

The E × B Electron Drift Instability (E × B EDI), also called Electron Cyclotron Drift Instability, has been observed in recent particle simulations of Hall thrusters and is a possible candidate to explain anomalous electron transport across the magnetic field in these devices. This instability is characterized by the development of an azimuthal wave with wavelength in the mm range and velocity on the order of the ion acoustic velocity, which enhances electron transport across the magnetic field. In this paper, we study the development and convection of the E × B EDI in the acceleration and near plume regions of a Hall thruster using a simplified 2D axial-azimuthal Particle-In-Cell simulation. The simulation is collisionless and the ionization profile is not-self-consistent but rather is given as an input parameter of the model. The aim is to study the development and properties of the instability for different values of the ionization rate (i.e., of the total ion production rate or current) and to compare the results with the theory. An important result is that the wavelength of the simulated azimuthal wave scales as the electron Debye length and that its frequency is on the order of the ion plasma frequency. This is consistent with the theory predicting destruction of electron cyclotron resonance of the E × B EDI in the non-linear regime resulting in the transition to an ion acoustic instability. The simulations also show that for plasma densities smaller than under nominal conditions of Hall thrusters the field fluctuations induced by the E × B EDI are no longer sufficient to significantly enhance electron transport across the magnetic field, and transit time instabilities develop in the axial direction. The conditions and results of the simulations are described in detail in this paper and they can serve as benchmarks for comparisons between different simulation codes. Such benchmarks would be very useful to study the role of numerical noise (numerical

Lithium position and occupancy fluctuations in a cathode during charge/discharge cycling of lithium-ion battery

International Nuclear Information System (INIS)

Sharma, N.; Yu, D.; Zhu, Y.; Wu, Y.; Peterson, V. K.

2012-01-01

Lithium-ion batteries are undergoing rapid development to meet the energy demands of the transportation and renewable energy-generation sectors. The capacity of a lithium-ion battery is dependent on the amount of lithium that can be reversibly incorporated into the cathode. Neutron diffraction provides greater sensitivity towards lithium relative to other diffraction techniques. In conjunction with the penetration depth afforded by neutron diffraction, the information concerning lithium gained in a neutron diffraction study allows commercial lithium-ion batteries to be explored with respect to the lithium content in the whole cathode. Furthermore, neutron diffraction instruments featuring area detectors that allow relatively fast acquisitions enable perturbations of lithium location and occupancy in the cathode during charge/discharge cycling to be determined in real time. Here, we present the time, current, and temperature dependent lithium transfer occurring within a cathode functioning under conventional charge-discharge cycling. The lithium location and content, oxygen positional parameter, and lattice parameter of the Li 1+y Mn 2 0 4 cathode are measured and linked to the battery's charge/discharge characteristics (performance). We determine that the lithium-transfer mechanism involves two crystallographic sites, and that the mechanism differs between discharge and charge, explaining the relative ease of discharging (compared with charging) this material. Furthermore, we find that the rate of change of the lattice is faster on charging than discharging, and is dependent on the lithium insertion/ extraction processes (e.g. dependent on how the site occupancies evolve). Using in situ neutron diffraction data the atomic-scale understanding of cathode functionality is revealed, representing detailed information that can be used to direct improvements in battery performance at both the practical and fundamental level.

High Precision Beam Diagnostics for Ion Thrusters

NARCIS (Netherlands)

Van Reijen, B.; Koch, N.; Lazurenko, A.; Weis, S.; Schirra, M.; Genovese, A.; Haderspeck, J.; Gill, E.K.A.

2011-01-01

The Thales diagnostic equipment for ion beam characterization consists of a gridded and single orifice retarding potential analyzer (RPA) and an energy selective mass spectrometer (ESMS). During the development phase of these sensors considerable effort was put into the removal of ion optical

Study on a negative hydrogen ion source with hot cathode arc discharge.

Science.gov (United States)

Lin, S H; Fang, X; Zhang, H J; Qian, C; Ma, B H; Wang, H; Li, X X; Zhang, X Z; Sun, L T; Zhang, Z M; Yuan, P; Zhao, H W

2014-02-01

A negative hydrogen (H(-)) ion source with hot cathode arc discharge was designed and fabricated as a primary injector for a 10 MeV PET cyclotron at IMP. 1 mA dc H(-) beam with ɛ N, RMS = 0.08 π mm mrad was extracted at 25 kV. Halbach hexapole was adopted to confine the plasma. The state of arc discharge, the parameters including filament current, arc current, gas pressure, plasma electrode bias, and the ratio of I(e(-))/I(H(-)) were experimentally studied. The discussion on the result, and opinions to improve the source were given.

Anode Fall Formation in a Hall Thruster

International Nuclear Information System (INIS)

Dorf, Leonid A.; Raitses, Yevgeny F.; Smirnov, Artem N.; Fisch, Nathaniel J.

2004-01-01

As was reported in our previous work, accurate, nondisturbing near-anode measurements of the plasma density, electron temperature, and plasma potential performed with biased and emissive probes allowed the first experimental identification of both electron-repelling (negative anode fall) and electron-attracting (positive anode fall) anode sheaths in Hall thrusters. An interesting new phenomenon revealed by the probe measurements is that the anode fall changes from positive to negative upon removal of the dielectric coating, which appears on the anode surface during the course of Hall thruster operation. As reported in the present work, energy dispersion spectroscopy analysis of the chemical composition of the anode dielectric coating indicates that the coating layer consists essentially of an oxide of the anode material (stainless steel). However, it is still unclear how oxygen gets into the thruster channel. Most importantly, possible mechanisms of anode fall formation in a Hall thruster with a clean and a coated anodes are analyzed in this work; practical implication of understanding the general structure of the electron-attracting anode sheath in the case of a coated anode is also discussed

Stability test and analysis of the Space Shuttle Primary Reaction Control Subsystem thruster

Science.gov (United States)

Applewhite, John; Hurlbert, Eric; Krohn, Douglas; Arndt, Scott; Clark, Robert

1992-01-01

The results are reported of a test program conducted on the Space Shuttle Primary Reaction Control Subsystem thruster in order to investigate the effects of trapped helium bubbles and saturated propellants on stability, determine if thruster-to-thruster stability variations are significant, and determine stability under STS-representative conditions. It is concluded that the thruster design is highly reliable in flight and that burn-through has not occurred. Significantly unstable thrusters are screened out, and wire wrap is found to protect against chamber burn-throughs and to provide a fail-safe thruster for this situation.

Spectrum Diagnosis for Fuchsia Plume of Hall Effect Thruster with Xenon as Propellant

International Nuclear Information System (INIS)

Yu Daren; Ding Jiapeng; Dai Jingmin

2006-01-01

The colour of the Hall effect thruster's plume is often light-green, and sometimes a fuchsia plume appears during experiments. Based on a spectrum and colour analysis, and a comparison with normal plumes, a conclusion is made that the density of the Xe ions and the temperature of electrons are low when the plume appears fuchsia. In this condition, most of the components of the plume are Xe atoms, and the ionization rate of the propellant is low

MPD thruster research issues, activities, strategies

Science.gov (United States)

1991-01-01

The following activities and plans in the MPD thruster development are summarized: (1) experimental and theoretical research (magnetic nozzles at present and high power levels, MPD thrusters with applied fields extending into the thrust chamber, and improved electrode performance); and (2) tools (MACH2 code for MPD and nozzle flow calculation, laser diagnostics and spectroscopy for non-intrusive measurements of flow conditions, and extension to higher power). National strategies are also outlined.

In-situ reactive glow discharge cleaning of NSLS distributed ion pumps

Energy Technology Data Exchange (ETDEWEB)

Johnson, E.D.; Chou, T.S.

1988-01-01

Based on our experience with the in-situ cleaning of optical systems by reactive r.f. glow discharges and the conditioning and preparation of distributed ion pump (DIP) elements, we have sought to develop strategies for recovering from severe vacuum accidents by restoring DIP elements of storage rings such as those at the NSLS in-situ. In this paper we will describe a series of experiments conducted in a test apparatus to condition a so called ''egg-crate'' DIP in-situ, (this older type element being common in older storage rings). A new untreated element which was unable to pump below 5x10 /sup /minus/8/ Torr in its initial condition was treated in oxygen and subsequent argon r.f. discharges utilizing the pump element as the discharge electrode producing a nitrogen pumping speed of 168 l/s at 2x10 /sup /minus/8/ Torr. A light bake at 75/degree/C increased this to nearly 500 l/s at 5x10 /sup /minus/8/ Torr. After exposure to atmosphere the speed was reduced to nil at these pressures but subsequently recovered, without bakeout, by glow discharge cleaning. 22 refs., 6 figs.

In-situ reactive glow discharge cleaning of NSLS distributed ion pumps

International Nuclear Information System (INIS)

Johnson, E.D.; Chou, T.S.

1988-01-01

Based on our experience with the in-situ cleaning of optical systems by reactive r.f. glow discharges and the conditioning and preparation of distributed ion pump (DIP) elements, we have sought to develop strategies for recovering from severe vacuum accidents by restoring DIP elements of storage rings such as those at the NSLS in-situ. In this paper we will describe a series of experiments conducted in a test apparatus to condition a so called ''egg-crate'' DIP in-situ, (this older type element being common in older storage rings). A new untreated element which was unable to pump below 5x10 /sup /minus/8/ Torr in its initial condition was treated in oxygen and subsequent argon r.f. discharges utilizing the pump element as the discharge electrode producing a nitrogen pumping speed of 168 l/s at 2x10 /sup /minus/8/ Torr. A light bake at 75/degree/C increased this to nearly 500 l/s at 5x10 /sup /minus/8/ Torr. After exposure to atmosphere the speed was reduced to nil at these pressures but subsequently recovered, without bakeout, by glow discharge cleaning. 22 refs., 6 figs

Influence of geometry of the discharge interval on distribution of ion and electron streams at surface of the Penning source cathode

International Nuclear Information System (INIS)

Egiazaryan, G.A.; Khachatrian, Zh.B.; Badalyan, E.S.; Ter-Gevorgyan, E.I.; Hovhannisyan, V.N.

2006-01-01

In the discharge of oscillating electrons, the mechanism of the processes, which controls the distribution of the ion and electron streams over the cathode surface, is investigated experimentally. The influence of the length of the discharge interval on value and distribution of the ion and electron streams is analyzed. The distribution both of ion and electron streams at the cathode surface is determined at different conditions of the discharge. It is shown that for given values of the anode diameter d a =31 mm and the gas pressure P=5x10 -5 Torr, the intensive stream of positive ions falls entirely on the cathode central area in the whole interval of the anode length variation (l a =1-11 cm). At the cathode, the ion current reaches the maximal value at a certain (optimal) value of the anode length that, in turn, depends on the anode voltage U a . The intensive stream of longitudinal electrons forms in the short anodes only (l a =2.5-3.5 cm) and depending on the choice of the discharge regime, may fall both on central and middle parts of the cathode

High Power MPD Thruster Development at the NASA Glenn Research Center

Science.gov (United States)

LaPointe, Michael R.; Mikellides, Pavlos G.; Reddy, Dhanireddy (Technical Monitor)

2001-01-01

Propulsion requirements for large platform orbit raising, cargo and piloted planetary missions, and robotic deep space exploration have rekindled interest in the development and deployment of high power electromagnetic thrusters. Magnetoplasmadynamic (MPD) thrusters can effectively process megawatts of power over a broad range of specific impulse values to meet these diverse in-space propulsion requirements. As NASA's lead center for electric propulsion, the Glenn Research Center has established an MW-class pulsed thruster test facility and is refurbishing a high-power steady-state facility to design, build, and test efficient gas-fed MPD thrusters. A complimentary numerical modeling effort based on the robust MACH2 code provides a well-balanced program of numerical analysis and experimental validation leading to improved high power MPD thruster performance. This paper reviews the current and planned experimental facilities and numerical modeling capabilities at the Glenn Research Center and outlines program plans for the development of new, efficient high power MPD thrusters.

Fault-Tolerant Region-Based Control of an Underwater Vehicle with Kinematically Redundant Thrusters

Directory of Open Access Journals (Sweden)

Zool H. Ismail

2014-01-01

Full Text Available This paper presents a new control approach for an underwater vehicle with a kinematically redundant thruster system. This control scheme is derived based on a fault-tolerant decomposition for thruster force allocation and a region control scheme for the tracking objective. Given a redundant thruster system, that is, six or more pairs of thrusters are used, the proposed redundancy resolution and region control scheme determine the number of thruster faults, as well as providing the reference thruster forces in order to keep the underwater vehicle within the desired region. The stability of the presented control law is proven in the sense of a Lyapunov function. Numerical simulations are performed with an omnidirectional underwater vehicle and the results of the proposed scheme illustrate the effectiveness in terms of optimizing the thruster forces.

Emissive Ion Thruster -EMIT, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — A propulsion system is proposed that is based on acceleration of ions emitted from a thin, solid-state electrochemical ceramic membrane. This technology would...

The FAST (FRC Acceleration Space Thruster) Experiment

Science.gov (United States)

Martin, Adam; Eskridge, R.; Lee, M.; Richeson, J.; Smith, J.; Thio, Y. C. F.; Slough, J.; Rodgers, Stephen L. (Technical Monitor)

2001-01-01

The Field Reverse Configuration (FRC) is a magnetized plasmoid that has been developed for use in magnetic confinement fusion. Several of its properties suggest that it may also be useful as a thruster for in-space propulsion. The FRC is a compact toroid that has only poloidal field, and is characterized by a high plasma beta = (P)/(B (sup 2) /2Mu0), the ratio of plasma pressure to magnetic field pressure, so that it makes efficient use of magnetic field to confine a plasma. In an FRC thruster, plasmoids would be repetitively formed and accelerated to high velocity; velocities of = 250 km/s (Isp = 25,000s) have already been achieved in fusion experiments. The FRC is inductively formed and accelerated, and so is not subject to the problem of electrode erosion. As the plasmoid may be accelerated over an extended length, it can in principle be made very efficient. And the achievable jet powers should be scalable to the MW range. A 10 kW thruster experiment - FAST (FRC Acceleration Space Thruster) has just started at the Marshall Space Flight Center. The design of FAST and the status of construction and operation will be presented.

Comparison of Medium Power Hall Effect Thruster Ion Acceleration for Krypton and Xenon Propellants

Science.gov (United States)

2016-09-14

Pumping is provided by four single-stage cryogenic panels (single-stage cold heads at 25 K) and one 50 cm two stage cryogenic pump (12 K). This vacuum...test chamber has a mea- sured pumping speed of 36 kL/s on xenon. The Hall thruster used in this study is a medium power laboratory Hall effect...The first compo- nent passes through a krypton opto-galvanic cell and is terminated by a beam dump . The opto-galvanic cell current is capacitively

Chemistry of sprite discharges through ion-neutral reactions

Science.gov (United States)

Hiraki, Y.; Kasai, Y.; Fukunishi, H.

2008-07-01

We estimate the concentration changes, caused by streamer discharge in sprites, of ozone and related minor species as odd nitrogen (NOx) and hydrogen (HOx) families in the upper stratosphere and mesosphere. The streamer has an intense electric field and high electron density at its head, where a large number of chemically-radical ions and atoms are produced through electron impact on neutral molecules. After its propagation, densities of minor species can be perturbed through ion-neutral chemical reactions initiated by the relaxation of these radical products. We evaluate the production rates of ions and atoms using an electron kinetics model and by assuming that the electric field and electron density are in the head region. We calculate the density variations mainly for NOx, Ox, and HOx species using a one-dimensional model of the neutral and ion composition of the middle atmosphere, including the effect of the sprite streamer. Results at the nighttime condition show that the densities of NO, O3, H, and OH increase suddenly through reactions triggered by the first atomic nitrogen and oxygen product, and electrons just after streamer initiation. It is shown that NO and NO2 still remain for 1 h by a certain order of increase with their source-sink balance, predominantly around 60 km; for other species, increases in O3, OH, HO2, and H2O2 still remain in the range of 40 70 km. From this affirmative result of long-time behavior previously not presented, we emphasize that sprites would have the power to impact local chemistry at night. We also discuss the consistency with previous theoretical and observational studies, along with future suggestions.

Particle-in-cell simulations of Hall plasma thrusters

Science.gov (United States)

Miranda, Rodrigo; Ferreira, Jose Leonardo; Martins, Alexandre

2016-07-01

Hall plasma thrusters can be modelled using particle-in-cell (PIC) simulations. In these simulations, the plasma is described by a set of equations which represent a coupled system of charged particles and electromagnetic fields. The fields are computed using a spatial grid (i.e., a discretization in space), whereas the particles can move continuously in space. Briefly, the particle and fields dynamics are computed as follows. First, forces due to electric and magnetic fields are employed to calculate the velocities and positions of particles. Next, the velocities and positions of particles are used to compute the charge and current densities at discrete positions in space. Finally, these densities are used to solve the electromagnetic field equations in the grid, which are interpolated at the position of the particles to obtain the acting forces, and restart this cycle. We will present numerical simulations using software for PIC simulations to study turbulence, wave and instabilities that arise in Hall plasma thrusters. We have sucessfully reproduced a numerical simulation of a SPT-100 Hall thruster using a two-dimensional (2D) model. In addition, we are developing a 2D model of a cylindrical Hall thruster. The results of these simulations will contribute to improve the performance of plasma thrusters to be used in Cubesats satellites currenty in development at the Plasma Laboratory at University of Brasília.

Investigation of multipactoring discharge in an H-type resonator of an ion linac

International Nuclear Information System (INIS)

Lobzev, L.D.; Mazalov, Yu.P.; Gusev, E.V.; Shulika, N.G.

1993-01-01

The study on the variation of parameters of an auto generator power supply of an ion linear accelerator which is under construction on the base of the H-type resonator with drift tube comb mounts has been made. The main attention was paid to investigating multipactor discharges. It is concluded, that presence of secondary autonomous self-oscillation circuit providing the field rise with the frequency close to operation frequency leads to the decrease of multipactor discharge intensity level, which is enough for the elimination of their harmful effect on accelerator steady-state performance. 8 refs., 3 figs

Advanced Propellants for Scalable, Multipurpose Electrospray Ion Thrusters

Data.gov (United States)

National Aeronautics and Space Administration — Ionic liquid ion sources (ILIS) have numerous applications in spacecraft propulsion and focused ion beam technologies. The Space Propulsion Lab at the Massachusetts...

Mission and System Advantages of Iodine Hall Thrusters

Science.gov (United States)

Dankanich, John W.; Szabo, James; Pote, Bruce; Oleson, Steve; Kamhawi, Hani

2014-01-01

The exploration of alternative propellants for Hall thrusters continues to be of interest to the community. Investments have been made and continue for the maturation of iodine based Hall thrusters. Iodine testing has shown comparable performance to xenon. However, iodine has a higher storage density and resulting higher ?V capability for volume constrained systems. Iodine's vapor pressure is low enough to permit low-pressure storage, but high enough to minimize potential adverse spacecraft-thruster interactions. The low vapor pressure also means that iodine does not condense inside the thruster at ordinary operating temperatures. Iodine is safe, it stores at sub-atmospheric pressure, and can be stored unregulated for years on end; whether on the ground or on orbit. Iodine fills a niche for both low power (10kW) electric propulsion regimes. A range of missions have been evaluated for direct comparison of Iodine and Xenon options. The results show advantages of iodine Hall systems for both small and microsatellite application and for very large exploration class missions.

Optimization of a coaxial electron cyclotron resonance plasma thruster with an analytical model

Energy Technology Data Exchange (ETDEWEB)

Cannat, F., E-mail: felix.cannat@onera.fr, E-mail: felix.cannat@gmail.com; Lafleur, T. [Physics and Instrumentation Department, Onera -The French Aerospace Lab, Palaiseau, Cedex 91123 (France); Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universites, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau (France); Jarrige, J.; Elias, P.-Q.; Packan, D. [Physics and Instrumentation Department, Onera -The French Aerospace Lab, Palaiseau, Cedex 91123 (France); Chabert, P. [Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universites, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau (France)

2015-05-15

A new cathodeless plasma thruster currently under development at Onera is presented and characterized experimentally and analytically. The coaxial thruster consists of a microwave antenna immersed in a magnetic field, which allows electron heating via cyclotron resonance. The magnetic field diverges at the thruster exit and forms a nozzle that accelerates the quasi-neutral plasma to generate a thrust. Different thruster configurations are tested, and in particular, the influence of the source diameter on the thruster performance is investigated. At microwave powers of about 30 W and a xenon flow rate of 0.1 mg/s (1 SCCM), a mass utilization of 60% and a thrust of 1 mN are estimated based on angular electrostatic probe measurements performed downstream of the thruster in the exhaust plume. Results are found to be in fair agreement with a recent analytical helicon thruster model that has been adapted for the coaxial geometry used here.

Density of uranium ions in the 4I0/sub 9/2/ ground state in a hollow-cathode type discharge

International Nuclear Information System (INIS)

Pianarosa, P.; Bouchard, P.; Saint-Dizier, J.P.; Gagne, J.M.

1983-01-01

A hollow-cathode type discharge cell as generator of uranium ions is investigated. The 4 I 0 /sub 9/2/ ground-state ion density has been obtained by absorption spectroscopy at 5493 and 4244 A. The absorption measurements have been performed using two identical hollow-cathode lamps: one acting as a light source, the other as a reservoir of free ions. Neon and xenon have been used as discharge sustaining gases. In our experimental conditions the measured ion ground-state density is of the order of 10 12 ions cm -3 . Absorption measurements performed at 5915 and 4246 A of U i give a density of the order of 10 12 atoms cm -3 . This latter value is in excellent agreement with a previously measured value obtained by laser-absorption spectroscopy

Influence of ion transport on discharge propagation of nanosecond dielectric barrier discharge plasma actuator

Science.gov (United States)

Hua, Weizhuo; Koji, Fukagata

2017-11-01

A numerical study has been conducted to understand the streamer formation and propagation of nanosecond pulsed surface dielectric barrier discharge of positive polarity. First we compared the result of different grid configuration to investigate the influence of x and y direction grid spacing on the streamer propagation. The streamer propagation is sensitive to y grid spacing especially at the dielectric surface. The streamer propagation velocity can reach 0.2 cm/ns when the voltage magnitude is 12 kV. A narrow gap was found between the streamer and dielectric barrier, where the plasma density is several orders of magnitude smaller than the streamer region. Analyses on the ion transport in the gap and streamer regions show the different ion transport mechanisms in the two different region. In the gap region, the diffusion of electron toward the dielectric layer decreases the seed electron in the beginning of voltage pulse, resulting that ionization avalanche does not occur. The streamer region is not significantly affected by the diffusion flux toward the dielectric layer, so that ionization avalanche takes place and leads to dramatic increase of plasma density.

Determination of soluble bromine in an extra-high-pressure mercury discharge lamp by sodium hydroxide decomposition-suppressed ion chromatography.

Science.gov (United States)

Mitsumata, Hiroshi; Mori, Toshio; Maeda, Tatsuo; Kita, Yoshiyuki; Kohatsu, Osamu

2006-02-01

We have established a simple method for assaying the quantity of soluble bromine in the discharge tubes of an extra-high-pressure mercury discharge lamp. Each discharge tube is destroyed in 5 ml of 10 mM sodium hydroxide, and the recovered sodium hydroxide solution is analyzed by suppressed-ion chromatography using gradient elution. We have clarified that this method can assay less than 1 microg of soluble bromine in a discharge tube.

Chemistry of sprite discharges through ion-neutral reactions

Directory of Open Access Journals (Sweden)

Y. Hiraki

2008-07-01

Full Text Available We estimate the concentration changes, caused by streamer discharge in sprites, of ozone and related minor species as odd nitrogen (NO_x and hydrogen (HO_x families in the upper stratosphere and mesosphere. The streamer has an intense electric field and high electron density at its head, where a large number of chemically-radical ions and atoms are produced through electron impact on neutral molecules. After its propagation, densities of minor species can be perturbed through ion-neutral chemical reactions initiated by the relaxation of these radical products. We evaluate the production rates of ions and atoms using an electron kinetics model and by assuming that the electric field and electron density are in the head region. We calculate the density variations mainly for NO_x, O_x, and HO_x species using a one-dimensional model of the neutral and ion composition of the middle atmosphere, including the effect of the sprite streamer. Results at the nighttime condition show that the densities of NO, O₃, H, and OH increase suddenly through reactions triggered by the first atomic nitrogen and oxygen product, and electrons just after streamer initiation. It is shown that NO and NO₂ still remain for 1 h by a certain order of increase with their source-sink balance, predominantly around 60 km; for other species, increases in O₃, OH, HO₂, and H₂O₂ still remain in the range of 40–70 km. From this affirmative result of long-time behavior previously not presented, we emphasize that sprites would have the power to impact local chemistry at night. We also discuss the consistency with previous theoretical and observational studies, along with future suggestions.

Sawtooth-free Ohmic discharges in ASDEX and the aspects of neoclassical ion transport

International Nuclear Information System (INIS)

Stroth, U.; Fussmann, G.; Krieger, K.; Mertens, V.; Wagner, F.; Bessenrodt-Weberpals, M.; Buechse, R.; Giannone, L.; Herrmann, H.; Simmet, E.; Steuer, K.H.

1991-05-01

Sawtooth-free Ohmic discharges can serve as a model case for a quiescent Tokamak plasma. We report on the properties and the global parameters of these discharges observed in ASDEX and make comments on the mechanism which seems to be responsible for the stabilization of the sawtooth instability. Stationary Ohmic discharge were used to investigate particle, impurity and energy transport in the absence of the sawtooth instability. Particular emphasis has been devoted to a comparison with the predictions of neoclassical theories. We find that the ion energy transport is on the level predicted by neoclassical theory and can explain particle and impurity transport with neoclassical inward drift velocities and diffusion coefficients with the same small anomalous contribution. In the central region of the plasma, where the power flux is low, very small values were found for the electron heat conductivity. (orig.)

Thermal Development Test of the NEXT PM1 Ion Engine

Science.gov (United States)

Anderson, John R.; Snyder, John S.; VanNoord, Jonathan L.; Soulas, George C.

2010-01-01

NASA's Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion propulsion system under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential mission destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic exploration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predicted hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

Negative ions formed in N₂/CH₄/Ar discharge – a simulation of Titan's atmosphere chemistry

OpenAIRE

Horvath, G.; Aranda-Gonzalvo, Y.; Mason, N. J.; Zahoran, M.; Matejcik, S.

2010-01-01

The formation of negative ions produced in a negative point-to-plane corona discharge fed by a Ar/N2//CH4/ gas mixture has been studied using mass spectrometry. The measurements were carried out in flowing regime at ambient temperature and a reduced pressure of 460 mbar. The CN ? anion has been found to be the most dominant negative ion in the discharge and is believed to be the precursor of heavier negative ions such as C3/N ? and C5/N ? . The most likely pathway for the formation of such mo...

Magnetic field deformation due to electron drift in a Hall thruster

Directory of Open Access Journals (Sweden)

Han Liang

2017-01-01

Full Text Available The strength and shape of the magnetic field are the core factors in the design of the Hall thruster. However, Hall current can affect the distribution of static magnetic field. In this paper, the Particle-In-Cell (PIC method is used to obtain the distribution of Hall current in the discharge channel. The Hall current is separated into a direct and an alternating part to calculate the induced magnetic field using Finite Element Method Magnetics (FEMM. The results show that the direct Hall current decreases the magnetic field strength in the acceleration region and also changes the shape of the magnetic field. The maximum reduction in radial magnetic field strength in the exit plane is 10.8 G for an anode flow rate of 15 mg/s and the maximum angle change of the magnetic field line is close to 3° in the acceleration region. The alternating Hall current induces an oscillating magnetic field in the whole discharge channel. The actual magnetic deformation is shown to contain these two parts.

Comment on 'Effects of Magnetic Field Gradient on Ion Beam Current in Cylindrical Hall Ion Source

International Nuclear Information System (INIS)

Raitses, Y.; Smirnov A.; Fisch, N.J.

2008-01-01

It is argued that the key difference of the cylindrical Hall thruster (CHT) as compared to the end-Hall ion source cannot be exclusively attributed to the magnetic field topology [Tang et al. J. Appl. Phys., 102, 123305 (2007)]. With a similar mirror-type topology, the CHT configuration provides the electric field with nearly equipotential magnetic field surfaces and a better suppression of the electron cross-field transport, as compared to both the end-Hall ion source and the cylindrical Hall ion source of Tang et al

Numerical simulation of SMART-1 Hall-thruster plasma interactions

NARCIS (Netherlands)

Tajmar, Martin; Sedmik, René; Scharlemann, Carsten

2009-01-01

SMART-1 has been the first European mission using a Hall thruster to reach the moon. An onboard plasma diagnostic package allowed a detailed characterization of the thruster exhaust plasma and its interactions with the spacecraft. Analysis of in-flight data revealed, amongst others, an unpredicted

Performance prediction of electrohydrodynamic thrusters by the perturbation method

International Nuclear Information System (INIS)

Shibata, H.; Watanabe, Y.; Suzuki, K.

2016-01-01

In this paper, we present a novel method for analyzing electrohydrodynamic (EHD) thrusters. The method is based on a perturbation technique applied to a set of drift-diffusion equations, similar to the one introduced in our previous study on estimating breakdown voltage. The thrust-to-current ratio is generalized to represent the performance of EHD thrusters. We have compared the thrust-to-current ratio obtained theoretically with that obtained from the proposed method under atmospheric air conditions, and we have obtained good quantitative agreement. Also, we have conducted a numerical simulation in more complex thruster geometries, such as the dual-stage thruster developed by Masuyama and Barrett [Proc. R. Soc. A 469, 20120623 (2013)]. We quantitatively clarify the fact that if the magnitude of a third electrode voltage is low, the effective gap distance shortens, whereas if the magnitude of the third electrode voltage is sufficiently high, the effective gap distance lengthens.

Measurement of electron- and ion beam energies and currents in a plasma focus discharge

International Nuclear Information System (INIS)

Yamamoto, Toshikazu; Kondoh, Yoshiomi; Shimoda, Katsuji; Hirano, Katsumi

1982-01-01

Measurements of energetic particle beams in a plsma focus with a Mather type device are presented. Rogowski coils are used for time-resolved measurement, and solid-state nuclear track detectors for time-integrated measurement of the beams. In the upstream direction with respect to the discharge current, only the electron beam with the maximum current of several kA was detected, which was approximately one percent of the discharge current. The electron energies of the beam were spread from 0.1 to 1 MeV. In the downstream direction, two successive emissions of ions were observed. The first emission had an extremely high energy of the order of some MeV and a low beam current of less than 10 A. The second emission, the main part of the ion beam, with energies of 100 - 800 keV, followed the first one with a time lag of several tens of nanoseconds, and the beam current reached several tens of amperes. (author)

Thermal stability of the krypton Hall effect thruster

Directory of Open Access Journals (Sweden)

Szelecka Agnieszka

2017-03-01

Full Text Available The Krypton Large IMpulse Thruster (KLIMT ESA/PECS project, which has been implemented in the Institute of Plasma Physics and Laser Microfusion (IPPLM and now is approaching its final phase, was aimed at incremental development of a ~500 W class Hall effect thruster (HET. Xenon, predominantly used as a propellant in the state-of-the-art HETs, is extremely expensive. Krypton has been considered as a cheaper alternative since more than fifteen years; however, to the best knowledge of the authors, there has not been a HET model especially designed for this noble gas. To address this issue, KLIMT has been geared towards operation primarily with krypton. During the project, three subsequent prototype versions of the thruster were designed, manufactured and tested, aimed at gradual improvement of each next exemplar. In the current paper, the heat loads in new engine have been discussed. It has been shown that thermal equilibrium of the thruster is gained within the safety limits of the materials used. Extensive testing with both gases was performed to compare KLIMT’s thermal behaviour when supplied with krypton and xenon propellants.

Parametric studies of the Hall Thruster at Soreq

International Nuclear Information System (INIS)

Ashkenazy, J.; Rattses, Y.; Appelbaum, G.

1997-01-01

An electric propulsion program was initiated at Soreq a few years ago, aiming at the research and development of advanced Hall thrusters for various space applications. The Hall 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/cm 2 ) can be obtained, since the acceleration mechanism is not limited by space charge effects. Such a device can be used as a small rocket engine onboard spacecraft with the advantage of a large jet velocity compared with conventional rocket engines (10,000-30,000 m/s vs. 2,000-4,800 m/s). An experimental Hall thruster was constructed at Soreq and operated under a broad range of operating conditions and under various configurational variations. Electrical, magnetic and plasma diagnostics, as well as 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. Representative results highlighting the major findings of the studies conducted so far are presented

Study on Endurance and Performance of Impregnated Ruthenium Catalyst for Thruster System.

Science.gov (United States)

Kim, Jincheol; Kim, Taegyu

2018-02-01

Performance and endurance of the Ru catalyst were studied for nitrous oxide monopropellant thruster system. The thermal decomposition of N2O requires a considerably high temperature, which make it difficult to be utilized as a thruster propellant, while the propellant decomposition temperature can be reduced by using the catalyst through the decomposition reaction with the propellant. However, the catalyst used for the thruster was frequently exposed to high temperature and high-pressure environment. Therefore, the state change of the catalyst according to the thruster operation was analyzed. Characterization of catalyst used in the operation condition of the thruster was performed using FE-SEM and EDS. As a result, performance degradation was occurred due to the volatilization of Ru catalyst and reduction of the specific surface area according to the phase change of Al2O3.

CO-AXIAL DISCHARGES

Science.gov (United States)

Luce, J.S.; Smith, L.P.

1960-11-22

A method and apparatus are given for producing coaxial arc discharges in an evacuated enclosure and within a strong, confining magnetic field. The arcs are maintained at a high potential difference. Electrons will diffuse to the more positive arc from the negative arc, and positive ions will diffuse from the more positive arc to the negative arc. Coaxial arc discharges have the advantage that ions which return to strike the positive arc discharge will lose no energy since they do not strike a solid wall or electrode. Those discharges are useful in confining an ionized plasma between the discharges, and have the advantage of preventing impurities from the walls of the enclosure from entering ihe plasma area because of the arc barrier set up bv the cylindrical outer arc.

Oxygen-Methane Thruster, Phase II

Data.gov (United States)

National Aeronautics and Space Administration — Two main innovations will be developed in the Phase II effort that are fundamentally associated with our gaseous oxygen/gaseous methane RCS thruster. The first...

Electrostatic Plasma Accelerator (EPA)

Science.gov (United States)

Brophy, John R.; Aston, Graeme

1995-01-01

The application of electric propulsion to communications satellites, however, has been limited to the use of hydrazine thrusters with electric heaters for thrust and specific impulse augmentation. These electrothermal thrusters operate at specific impulse levels of approximately 300 s with heater powers of about 500 W. Low power arcjets (1-3 kW) are currently being investigated as a way to increase specific impulse levels to approximately 500 s. Ion propulsion systems can easily produce specific impulses of 3000 s or greater, but have yet to be applied to communications satellites. The reasons most often given for not using ion propulsion systems are their high level of overall complexity, low thrust with long burn times, and the difficulty of integrating the propulsion system into existing commercial spacecraft busses. The Electrostatic Plasma Accelerator (EPA) is a thruster concept which promises specific impulse levels between low power arcjets and those of the ion engine while retaining the relative simplicity of the arcjet. The EPA thruster produces thrust through the electrostatic acceleration of a moderately dense plasma. No accelerating electrodes are used and the specific impulse is a direct function of the applied discharge voltage and the propellant atomic mass.

Current-Induced Membrane Discharge

DEFF Research Database (Denmark)

Andersen, Mathias Bækbo; van Soestbergen, M.; Mani, A.

2012-01-01

. Salt depletion leads to a large electric field resulting in a local pH shift within the membrane with the effect that the membrane discharges and loses its ion selectivity. Since salt co-ions, H+ ions, and OH- ions contribute to OLC, CIMD interferes with electrodialysis (salt counterion removal...... neglects chemical effects and remains to be quantitatively tested. Here, we show that charge regulation and water self-ionization can lead to OLC by "current-induced membrane discharge'' (CIMD), even in the absence of fluid flow, in ion-exchange membranes much thicker than the local Debye screening length...

Argon discharge characteristics in cold cathode penning ion source. Vol. 2

Energy Technology Data Exchange (ETDEWEB)

Abd El-Baki, M M; Abd El-Rahman, M M; Basal, N I [Ion Sources and Accelerators Department, Nuclear Research Center, Atomic energy Authority, Cairo, (Egypt)

1996-03-01

This study includes the production of argon discharge inside cold cathode penning ion source with axial d.c. extraction. The arc characteristics are investigated under the influence of the discharge parameters such as the pressure, axial magnetic field. At zero magnetic field and pressure 4.2 x 10{sup -4} torr, the arc voltage which is needed for arc initiation is relatively large V{sub arc} = 430 V, and I{sub arc} = 0.3 A. The application of the magnetic field helps the appearance of argon arc at lower voltage, e.g. at I{sub B} = 0.8 A, the arc voltage V{sub arc} = 320 V, and I{sub arc} = 0.3 A. It is found that the arc current increase with the increase of pressure, i.e. the increase of gas flow inside the source, while the arc voltage decreases. 7 fig.

Characteristics of a non-volatile liquid propellant in liquid-fed ablative pulsed plasma thrusters

Science.gov (United States)

Ling, William Yeong Liang; Schönherr, Tony; Koizumi, Hiroyuki

2017-02-01

In the past several decades, the use of electric propulsion in spacecraft has experienced tremendous growth. With the increasing adoption of small satellites in the kilogram range, suitable propulsion systems will be necessary in the near future. Pulsed plasma thrusters (PPTs) were the first form of electric propulsion to be deployed in orbit, and are highly suitable for small satellites due to their inherent simplicity. However, their lifetime is limited by disadvantages such as carbon deposition leading to thruster failure, and complicated feeding systems required due to the conventional use of solid propellants (usually polytetrafluoroethylene (PTFE)). A promising alternative to solid propellants has recently emerged in the form of non-volatile liquids that are stable in vacuum. This study presents a broad comparison of the non-volatile liquid perfluoropolyether (PFPE) and solid PTFE as propellants on a PPT with a common design base. We show that liquid PFPE can be successfully used as a propellant, and exhibits similar plasma discharge properties to conventional solid PTFE, but with a mass bit that is an order of magnitude higher for an identical ablation area. We also demonstrate that the liquid PFPE propellant has exceptional resistance to carbon deposition, completely negating one of the major causes of thruster failure, while solid PTFE exhibited considerable carbon build-up. Energy dispersive X-ray spectroscopy was used to examine the elemental compositions of the surface deposition on the electrodes and the ablation area of the propellant (or PFPE encapsulator). The results show that based on its physical characteristics and behavior, non-volatile liquid PFPE is an extremely promising propellant for use in PPTs, with an extensive scope available for future research and development.

Simulations of momentum transfer process between solar wind plasma and bias voltage tethers of electric sail thruster

Science.gov (United States)

Xia, Guangqing; Han, Yajie; Chen, Liuwei; Wei, Yanming; Yu, Yang; Chen, Maolin

2018-06-01

The interaction between the solar wind plasma and the bias voltage of long tethers is the basic mechanism of the electric sail thruster. The momentum transfer process between the solar wind plasma and electric tethers was investigated using a 2D full particle PIC method. The coupled electric field distribution and deflected ion trajectory under different bias voltages were compared, and the influence of bias voltage on momentum transfer process was analyzed. The results show that the high potential of the bias voltage of long tethers will slow down, stagnate, reflect and deflect a large number of ions, so that ion cavities are formed in the vicinity of the tether, and the ions will transmit the axial momentum to the sail tethers to produce the thrust. Compared to the singe tether, double tethers show a better thrust performance.

Particle-in-cell simulation for the effect of segmented electrodes near the exit of an aton-type Hall thruster on ion focusing acceleration

Energy Technology Data Exchange (ETDEWEB)

Yu, D.R.; Qing, S.W.; Liu, H.; Li, H. [Lab. of Plasma Propulsion, Harbin Institute of Technology (China)

2011-12-15

The effect of floating conductive electrodes near the channel exit of an Aton-type Hall thruster on ion focusing acceleration is studied by simulating the two-dimensional plasma flow with a fully kinetic Particle-in-Cell method for the gas flow rate j{sub a} ranged in 1{proportional_to}3 mg/s. Numerical results show that low-emissive electrodes can reduce plume divergence if the electrode length is less than 2 mm due to the low secondary electron emissive characteristic, but widen plume in all the gas flow rate range if the electrode length is greater than 2mm since the conductive property of segmented electrodes trends to make equipotential lines convex toward channel exit and is even parallel to the wall surface in the near-wall region. Further investigation predicts that the combination of high emissive dielectric wall and segmented low-emissive dielectric wall is a promising way to reduce plume divergence (copyright 2011 WILEY-VCH Verlag GmbH and Co. KGaA, Weinheim) (orig.)

Multi-Axis Thrust Measurements of the EO-1 Pulsed Plasma Thruster

Science.gov (United States)

Arrington, Lynn A.; Haag, Thomas W.

1999-01-01

Pulsed plasma thrusters are low thrust propulsive devices which have a high specific impulse at low power. A pulsed plasma thruster is currently scheduled to fly as an experiment on NASA's Earth Observing-1 satellite mission. The pulsed plasma thruster will be used to replace one of the reaction wheels. As part of the qualification testing of the thruster it is necessary to determine the nominal thrust as a function of charge energy. These data will be used to determine control algorithms. Testing was first completed on a breadboard pulsed plasma thruster to determine nominal or primary axis thrust and associated propellant mass consumption as a function of energy and then later to determine if any significant off-axis thrust component existed. On conclusion that there was a significant off-axis thrust component with the bread-board in the direction of the anode electrode, the test matrix was expanded on the flight hardware to include thrust measurements along all three orthogonal axes. Similar off-axis components were found with the flight unit.

Development of HAN-based Liquid Propellant Thruster

Science.gov (United States)

Hisatsune, K.; Izumi, J.; Tsutaya, H.; Furukawa, K.

2004-10-01

Many of propellants that are applied to the conventional spacecraft propulsion system are toxic propellants. Because of its toxicity, considering the environmental pollution or safety on handling, it will be necessary to apply the "green" propellant to the spacecraft propulsion system. The purpose of this study is to apply HAN based liquid propellant (LP1846) to mono propellant thruster. Compared to the hydrazine that is used in conventional mono propellant thruster, HAN based propellant is not only lower toxic but also can obtain higher specific impulse. Moreover, HAN based propellant can be decomposed by the catalyst. It means there are the possibility of applying to the mono propellant thruster that can leads to the high reliability of the propulsion system.[1],[2] However, there are two technical subjects, to apply HAN based propellant to the mono propellant thruster. One is the high combustion temperature. The catalyst will be damaged under high temperature condition. The other is the low catalytic activity. It is the serious problem on application of HAN based propellant to the mono propellant thruster that is used for attitude control of spacecraft. To improve the catalytic activity of HAN based propellant, it is necessary to screen the best catalyst for HAN based propellant. The adsorption analysis is conducted by Monte Carlo Simulation to screen the catalyst metal for HAN and TEAN. The result of analysis shows the Iridium is the best catalyst metal for HAN and TEAN. Iridium is the catalyst metal that is used at conventional mono propellant thruster catalyst Shell405. Then, to confirm the result of analysis, the reaction test about catalyst is conducted. The result of this test is the same as the result of adsorption analysis. That means the adsorption analysis is effective in screening the catalyst metal. At the evaluating test, the various types of carrier of catalyst are also compared to Shell 405 to improve catalytic activity. The test result shows the

The Plasmoid Thruster Experiment (PTX)

Science.gov (United States)

Eskridge, Richard; Martin, Adam; Koelfgen, Syri; Lee, Mike; Smith, James W.

2003-01-01

A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are categorized according to the relative strength of the poloidal and toroidal magnetic field (B(phi), and B(tau), respectively). An object with B(phi)/B(tau) >> 1 is classified as a Field Reverse Configuration (FRC); if B(phi) = B(tau), it is called a Spheromak. There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A thruster based on this concept would operate by repetitively producing plasmoids and ejecting them from the device at high velocity. The plasmoid is formed inside of a single turn conical theta-pinch coil; as this process is inductive, there are no life-limiting electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s (l), and calculations indicate that velocities in excess of 100 km/s are possible. A thruster based on this concept would be capable of producing an I(sp) in the range of 5,000 - 10,OOO s, with thrust densities of order 10(exp 5) N/m(exp 2). The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable to higher power. The purpose of this experiment is to determine the feasibility of this plasma propulsion concept. To accomplish this, it will be necessary to determine: a.) specific impulse and thrust, b.) efficiency and mass utilization, c.) which type of plasmoid (FRC-like or Spheromak-like) gives the best performance, and d.) the characteristics required of actual thruster components (i.e., switch and capacitor technology). The plasmoid mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras, and an interferometer. Simulations of the plasmoid thruster using MOQUI, a time dependent MHD code, will be carried out concurrently with experimental testing. The PTX

DESIGN AND DEVELOPMENT OF AUTO DEPTH CONTROL OF REMOTELY OPERATED VEHICLE USING THRUSTER SYSTEM

Directory of Open Access Journals (Sweden)

F.A. Ali

2014-12-01

Full Text Available Remotely Operated Vehicles are underwater robots designed specifically for surveillance, monitoring and collecting data for underwater activities. In the underwater vehicle industries, the thruster is an important part in controlling the direction, depth and speed of the ROV. However, there are some ROVs that cannot be maintained at the specified depth for a long time because of disturbance. This paper proposes an auto depth control using a thruster system. A prototype of a thruster with an auto depth control is developed and attached to the previously fabricated UTeM ROV. This paper presents the operation of auto depth control as well as thrusters for submerging and emerging purposes and maintaining the specified depth. The thruster system utilizes a microcontroller as its brain, a piezoresistive strain gauge pressure sensor and a DC brushless motor to run the propeller. Performance analysis of the auto depth control system is conducted to identify the sensitivity of the pressure sensor, and the accuracy and stability of the system. The results show that the thruster system performs well in maintaining a specified depth as well as stabilizing itself when a disturbanceoccurs even with a simple proportional controller used to control the thruster, where the thruster is an important component of the ROV.

Effects of cusped field thruster on the performance of drag-free control system

Science.gov (United States)

Cui, K.; Liu, H.; Jiang, W. J.; Sun, Q. Q.; Hu, P.; Yu, D. R.

2018-03-01

With increased measurement tasks of space science, more requirements for the spacecraft environment have been put forward. Those tasks (e.g. the measurement of Earth's steady state gravity field anomalies) lead to the desire for developing drag-free control. Higher requirements for the thruster performance are made due to the demand for the drag-free control system and real-time compensation for non-conservative forces. Those requirements for the propulsion system include wide continuous throttling ability, high resolution, rapid response, low noise and so on. As a promising candidate, the cusped field thruster has features such as the high working stability, the low erosion rate, a long lifetime and the simple structure, so that it is chosen as the thruster to be discussed in this paper. Firstly, the performance of a new cusped field thruster is tested and analyzed. Then a drag-free control scheme based on the cusped field thruster is designed to evaluate the performance of this thruster. Subsequently, the effects of the thrust resolution, transient response time and thrust uncertainty on the controller are calculated respectively. Finally, the performance of closed-loop system is analyzed, and the simulation results verify the feasibility of applying cusped field thruster to drag-free flight in the space science measurement tasks.

Double layers in a modestly collisional electronegative discharge

CERN Document Server

Sheridan, T E

1999-01-01

The effect of ion-neutral collisions on the structure and ion flux emanating from a steady-state, planar discharge with two negative components is investigated. The positive ion component is modelled as a cold fluid subject to constant-mobility collisions, while the electrons and negative ions obey Boltzmann relations. The model includes the collisionless limit. When the negative ions are sufficiently cold three types of discharge structures are found. For small negative ion concentrations or high collisionality, the discharge is 'stratified', with an electronegative core and an electropositive edge. For the opposite conditions, the discharge is 'uniform' with the negative ion density remaining significant at the edge of the plasma. Between these cases lies the special case of a double-layer-stratified discharge, where quasi-neutrality is violated at the edge of the electronegative core. Double-layer-stratified solutions are robust in that they persist for moderate collisionality. Numerical solutions for fini...

Electric-discharge-pumped nitrogen ion laser

Science.gov (United States)

Laudenslager, J. B.; Pacala, T. J.; Wittig, C.

1976-01-01

The routine operation is described of an N2(+) laser oscillating on the first negative band system of N2(+) which is produced in a preionized transverse discharge device. The discharge design incorporates features which favor the efficient production of the excitation transfer reaction of He2(+) with N2. A capacitive discharge switched by means of a high-current grounded grid thyratron is used to meet the design requirement of a volumetric discharge in high-pressure gas mixtures where the electric discharge need not have an ultrafast rise time (greater than 10 nsec) but should be capable of transferring large quantities of stored electric energy to the gas. A peak power of 180 kW in an 8-nsec laser pulse was obtained with a 0.1% mixture of N2 in helium at a total pressure of 3 atm. The most intense laser oscillations were observed on the (0,1) vibrational transition at 427.8 microns.

Use of Cumulative Degradation Factor Prediction and Life Test Result of the Thruster Gimbal Assembly Actuator for the Dawn Flight Project

Science.gov (United States)

Lo, C. John; Brophy, John R.; Etters, M. Andy; Ramesham, Rajeshuni; Jones, William R., Jr.; Jansen, Mark J.

2009-01-01

The Dawn Ion Propulsion System is the ninth project in NASA s Discovery Program. The Dawn spacecraft is being developed to enable the scientific investigation of the two heaviest main-belt asteroids, Vesta and Ceres. Dawn is the first mission to orbit two extraterrestrial bodies, and the first to orbit a main-belt asteroid. The mission is enabled by the onboard Ion Propulsion System (IPS) to provide the post-launch delta-V. The three Ion Engines of the IPS are mounted on Thruster Gimbal Assembly (TGA), with only one engine operating at a time for this 10-year mission. The three TGAs weigh 14.6 kg.

Combined tunable diode laser absorption spectroscopy and monochromatic radiation thermometry in ammonium dinitramide-based thruster

Science.gov (United States)

Zeng, Hui; Ou, Dongbin; Chen, Lianzhong; Li, Fei; Yu, Xilong

2018-02-01

Nonintrusive temperature measurements for a real ammonium dinitramide (ADN)-based thruster by using tunable diode laser absorption spectroscopy and monochromatic radiation thermometry are proposed. The ADN-based thruster represents a promising future space propulsion employing green, nontoxic propellant. Temperature measurements in the chamber enable quantitative thermal analysis for the thruster, providing access to evaluate thermal properties of the thruster and optimize thruster design. A laser-based sensor measures temperature of combustion gas in the chamber, while a monochromatic thermometry system based on thermal radiation is utilized to monitor inner wall temperature in the chamber. Additional temperature measurements of the outer wall temperature are conducted on the injector, catalyst bed, and combustion chamber of the thruster by using thermocouple, respectively. An experimental ADN thruster is redesigned with optimizing catalyst bed length of 14 mm and steady-state firing tests are conducted under various feed pressures over the range from 5 to 12 bar at a typical ignition temperature of 200°C. A threshold of feed pressure higher than 8 bar is required for the thruster's normal operation and upstream movement of the heat release zone is revealed in the combustion chamber out of temperature evolution in the chamber.

The development of the micro-solid propellant thruster array with improved repeatability

International Nuclear Information System (INIS)

Seo, Daeban; Kwon, Sejin; Lee, Jongkwang

2012-01-01

This paper presents the development of a micro-solid propellant thruster array with improved repeatability. The repeatability and low performance variation of each thruster unit with a high ignition success rate is essential in micro-solid propellant thruster array. To date, the study on the improvement of the repeatability has not yet been reported. As the first step for this study, we propose a new type of micro igniter, using a glass wafer called the heater-contact micro igniter. This igniter is also designed to improve the ignition characteristics of a glass-based micro igniter. The prototype of the igniter array is designed and fabricated to establish its fabrication process and to conduct its performance evaluation. Through the firing test, the performance of the heater-contact micro igniter is verified. The 5 × 5 sized micro-solid propellant thruster array is designed and fabricated applying the developed heater-contact igniter. The measured average thrust of each thruster unit is 2.542 N, and calculated standard deviation is 0.369 N. The calculated average total impulse and its standard deviation are 0.182 and 0.04 mNs, respectively. Based on these results, the improvement of repeatability is verified. Finally, the ignition control system of the micro-thruster array is developed. (paper)

Measurements of actinometry and ions energy in a microwave discharge

International Nuclear Information System (INIS)

Becerril, F.; Camps, E.; Villagran, M.; Muhl, S.

1998-01-01

In the present work it is showed the implementation of the plasma diagnostic technique through actinometry which allows to determine the absolute density of excited species. It is showed the range of the technique application, for the case of N 2 -H 2 mixtures plasmas used for the metals nitridation. The effects of magnetic field and the work pressure over ions energy were determined, using a Faraday cup type energy analyser. The results showed that in our device it is possible to vary such energy in a range between 10-45 eV, which amplify the range of applications perceptibly in comparison with another type of discharges. (Author)

Carbon Nanotube Based Electric Propulsion Thruster with Low Power Consumption, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — This SBIR project is to develop field emission electric propulsion (FEEP) thruster using carbon nanotubes (CNT) integrated anode. FEEP thrusters have gained...

Mathematical Modeling of Liquid-fed Pulsed Plasma Thruster

Directory of Open Access Journals (Sweden)

Kaartikey Misra

2018-01-01

Full Text Available Liquid propellants are fast becoming attractive for pulsed plasma thrusters due to their high efficiency and low contamination issues. However, the complete plasma interaction and acceleration processes are still not very clear. Present paper develops a multi-layer numerical model for liquid propellant PPTs (pulsed plasma thrusters. The model is based on a quasi-steady flow assumption. The model proposes a possible acceleration mechanism for liquid-fed pulsed plasma thrusters and accurately predicts the propellant utilization capabilities and estimations for the fraction of propellant gas that is completely ionized and accelerated to high exit velocities. Validation of the numerical model and the assumptions on which the model is based on is achieved by comparing the experimental results and the simulation results for two different liquid-fed thrusters developed at the University of Tokyo. Simulation results shows that up-to 50 % of liquid propellant injected is completely ionized and accelerated to high exit velocities (>50 Km/s, whereas, neutral gas contribute to only 7 % of the total specific impulse and accelerated to low exit velocity (<4 Km/s. The model shows an accuracy up-to 92 % . Optimization methods are briefly discussed to ensure efficient propellant utilization and performance. The model acts as a tool to understand the background physics and to optimize the performance for liquid-fed PPTs.

Electronegative Gas Thruster - Direct Thrust Measurement

Data.gov (United States)

National Aeronautics and Space Administration — This effort is an international collaboration and academic partnership to mature an innovative electric propulsion (EP) thruster concept to TRL 3 through direct...

Facility Effect Characterization Test of NASA's HERMeS Hall Thruster

Science.gov (United States)

Huang, Wensheng; Kamhawi, Hani; Haag, Thomas W.; Ortega, Alejandro Lopez; Mikellides, Ioannis G.

2016-01-01

A test to characterize the effect of varying background pressure on NASA's 12.5-kW Hall Effect Rocket with Magnetic Shielding had being completed. This thruster is the baseline propulsion system for the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). Potential differences in thruster performance and oscillation characteristics when in ground facilities versus on-orbit are considered a primary risk for the propulsion system of the Asteroid Redirect Robotic Mission, which is a candidate for SEP TDM. The first primary objective of this test was to demonstrate that the tools being developed to predict the zero-background-pressure behavior of the thruster can provide self-consistent results. The second primary objective of this test was to provide data for refining a physics-based model of the thruster plume that will be used in spacecraft interaction studies. Diagnostics deployed included a thrust stand, Faraday probe, Langmuir probe, retarding potential analyzer, Wien filter spectrometer, and high-speed camera. From the data, a physics-based plume model was refined. Comparisons of empirical data to modeling results are shown.

Overview of NASA GRCs Green Propellant Infusion Mission Thruster Testing and Plume Diagnostics

Science.gov (United States)

Deans, Matthew C.; Reed, Brian D.; Yim, John T.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; McLean, Christopher H.

2014-01-01

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The models describe the pressure, temperature, density, Mach number, and species concentration of the AF-M315E thruster exhaust plumes. The models are being used to assess the impingement effects of the AF-M315E thrusters on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters will be tested in a small rocket, altitude facility at NASA GRC. The GRC thruster testing will be conducted at duty cycles representatives of the planned GPIM maneuvers. A suite of laser-based diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, Schlieren imaging, and physical probes will be used to acquire plume measurements of AFM315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

Continuous Wheel Momentum Dumping Using Magnetic Torquers and Thrusters

Science.gov (United States)

Oh, Hwa-Suk; Choi, Wan-Sik; Eun, Jong-Won

1996-12-01

Two momentum management schemes using magnetic torquers and thrusters are sug-gested. The stability of the momentum dumping logic is proved at a general attitude equilibrium. Both momentum dumping control laws are implemented with Pulse-Width- Pulse-Frequency Modulated on-off control, and shown working equally well with the original continuous and variable strength control law. Thrusters are assummed to be asymmetrically configured as a contingency case. Each thruster is fired following separated control laws rather than paired thrusting. Null torque thrusting control is added on the thrust control calculated from the momentum control law for the gener-ation of positive thrusting force. Both magnetic and thrusting control laws guarantee the momentum dumping, however, the wheel inner loop control is needed for the "wheel speed" dumping, The control laws are simulated on the KOrea Multi-Purpose SATellite (KOMPSAT) model.

Coincident ion acceleration and electron extraction for space propulsion using the self-bias formed on a set of RF biased grids bounding a plasma source

International Nuclear Information System (INIS)

Rafalskyi, D; Aanesland, A

2014-01-01

We propose an alternative method to accelerate ions in classical gridded ion thrusters and ion sources such that co-extracted electrons from the source may provide beam space charge neutralization. In this way there is no need for an additional electron neutralizer. The method consists of applying RF voltage to a two-grid acceleration system via a blocking capacitor. Due to the unequal effective area of the two grids in contact with the plasma, a dc self-bias is formed, rectifying the applied RF voltage. As a result, ions are continuously accelerated within the grid system while electrons are emitted in brief instants within the RF period when the RF space charge sheath collapses. This paper presents the first experimental results and a proof-of-principle. Experiments are carried out using the Neptune thruster prototype which is a gridded Inductively Coupled Plasma (ICP) source operated at 4 MHz, attached to a larger beam propagation chamber. The RF power supply is used both for the ICP discharge (plasma generation) and powering the acceleration grids via a capacitor for ion acceleration and electron extraction without any dc power supplies. The ion and electron energies, particle flux and densities are measured using retarding field energy analyzers (RFEA), Langmuir probes and a large beam target. The system operates in Argon and N 2 . The dc self-bias is found to be generated within the gridded extraction system in all the range of operating conditions. Broad quasi-neutral ion-electron beams are measured in the downstream chamber with energies up to 400 eV. The beams from the RF acceleration method are compared with classical dc acceleration with an additional external electron neutralizer. It is found that the two acceleration techniques provide similar performance, but the ion energy distribution function from RF acceleration is broader, while the floating potential of the beam is lower than for the dc accelerated beam. (paper)

Modeling of physical processes in radio-frequency plasma thrusters

OpenAIRE

Tian, Bin

2017-01-01

This Thesis presents an investigation of the plasma-wave interaction in Helicon Plasma Thrusters (HPT). The HPT is a new concept of electric space propulsion, which generates plasmas with RF heating and provides thrust by the electrodeless acceleration of plasmas in a magnetic nozzle. An in-depth and extensive literature review of the state of the art of the models and experiments of plasma-wave interaction in helicon plasma sources and thrusters is carried out. Then, a theoret...

Power balance in ELMO Bumpy Torus: bulk electrons and ions in a 37 kW discharge

International Nuclear Information System (INIS)

McNeill, D.H.

1985-10-01

The power balance of the bulk electrons and ions in discharges with 37 kW of applied microwave power in the ELMO Bumpy Torus (EBT) is examined in a zero-dimensional model using data on the intensity and linewidth of the molecular and atomic hydrogen emission. At least 60% of the applied power is ultimately dissipated by processes involving the neutral particles, including dissociation of molecules, ionization of and radiation from atoms, and heating of cold electrons produced during atomic ionization. The molecular influx rate and the density of atoms are used independently to determine the bulk electron particle confinement time, and an upper bound estimate is made of the diffusional power loss from the bulk plasma electrons. Parameters derived from the basic spectroscopic data presented in this paper include the neutral atom density 2 - 5x10 10 cm -3 , incident molecular flux 3 - 5x10 15 cm -2 s -1 , bulk ion temperature approx. =3 eV, and particle confinement time <1.1 ms. The bulk electron energy confinement time is 0.7 ms or less in the standard operating regime. Published data on the nonthermal electron and ion populations in the plasma are used to evaluate approximately the overall energy flow in the discharge. 54 refs

Pulsed ion beam-assisted carburizing of titanium in methane discharge

International Nuclear Information System (INIS)

Shafiq, M.; Shahzad, K.; Zakaullah, M.; Hassan, M.; Qayyum, A.; Ahmad, S.; Rawat, R. S.

2010-01-01

The carburizing of titanium (Ti) is accomplished by utilizing energetic ion pulses of a 1.5 kJ Mather type dense plasma focus (DPF) device operated in methane discharge. X-ray diffraction (XRD) analysis confirms the deposition of polycrystalline titanium carbide (TiC). The samples carburized at lower axial and angular positions show an improved texture for a typical (200)TiC plane. The Williamson–Hall method is employed to estimate average crystallite size and microstrains in the carburized Ti surface. Crystallite size is found to vary from ∼ 50 to 100 nm, depending on the deposition parameters. Microstrains vary with the sample position and hence ion flux, and are converted from tensile to compressive by increasing the flux. The carburizing of Ti is confirmed by two major doublets extending from 300 to 390 cm −1 and from 560 to 620 cm −1 corresponding to acoustic and optical active modes in Raman spectra, respectively. Analyses by scanning electron microscopy/energy dispersive x-ray spectroscopy (SEM/EDS) have provided qualitative and quantitative profiles of the carburized surface. The Vickers microhardness of Ti is significantly improved after carburizing. (nuclear physics)

Characteristics of the LeRC/Hughes J-series 30-cm engineering model thruster

Science.gov (United States)

Collett, C. R.; Poeschel, R. L.; Kami, S.

1981-01-01

As a consequence of endurance and structural tests performed on 900-series engineering model thrusters (EMT), several modifications in design were found to be necessary for achieving performance goals. The modified thruster is known as the J-series EMT. The most important of the design modifications affect the accelerator grid, gimbal mount, cathode polepiece, and wiring harness. The paper discusses the design modifications incorporated, the condition(s) they corrected, and the characteristics of the modified thruster.

Chaos in gas discharges

International Nuclear Information System (INIS)

Piel, A.

1993-01-01

Many gas discharges exhibit natural oscillations which undergo a transition from regular to chaotic behavior by changing an experimental parameter or by applying external modulation. Besides several isolated investigations, two classes of discharge phenomena have been studied in more detail: ionization waves in medium pressure discharges and potential relaxation oscillations in filament cathode discharges at very low pressure. The latter phenomenon will be discussed by comparing experimental results from different discharge arrangements with particle-in-cell simulations and with a model based on the van-der-Pol equation. The filament cathode discharge has two stable modes of operation: the low current anode-glow-mode and the high current temperature-limited-mode, which form the hysteresis curve in the I(U) characteristics. Close to the hysteresis point of the AGM periodic relaxation oscillations occur. The authors demonstrate that the AGM can be understood by ion production in the anode layer, stopping of ions by charge exchange, and trapping in the virtual cathode around the filament. The relaxation oscillations consist of a slow filling phase and a rapid phase that invokes formation of an unstable double-layer, current-spiking, and ion depletion from the cathodic plasma. The relaxation oscillations can be mode-locked by external modulation. Inside a mode-locked state, a period doubling cascade is observed at high modulation degree

Thruster allocation for dynamical positioning

NARCIS (Netherlands)

Poppe, K.; van den Berg, J.B.; Blank, E.; Archer, C.; Redeker, M.; Kutter, M.; Hemker, P.

2010-01-01

Positioning a vessel at a fixed position in deep water is of great importance when working offshore. In recent years a Dynamical Positioning (DP) system was developed at Marin [2]. After the measurement of the current position and external forces (like waves, wind etc.), each thruster of the vessel

Micropulsed Plasma Thrusters for Attitude Control of a Low-Earth-Orbiting CubeSat

Science.gov (United States)

Gatsonis, Nikolaos A.; Lu, Ye; Blandino, John; Demetriou, Michael A.; Paschalidis, Nicholas

2016-01-01

This study presents a 3-Unit CubeSat design with commercial-off-the-shelf hardware, Teflon-fueled micropulsed plasma thrusters, and an attitude determination and control approach. The micropulsed plasma thruster is sized by the impulse bit and pulse frequency required for continuous compensation of expected maximum disturbance torques at altitudes between 400 and 1000 km, as well as to perform stabilization of up to 20 deg /s and slew maneuvers of up to 180 deg. The study involves realistic power constraints anticipated on the 3-Unit CubeSat. Attitude estimation is implemented using the q method for static attitude determination of the quaternion using pairs of the spacecraft-sun and magnetic-field vectors. The quaternion estimate and the gyroscope measurements are used with an extended Kalman filter to obtain the attitude estimates. Proportional-derivative control algorithms use the static attitude estimates in order to calculate the torque required to compensate for the disturbance torques and to achieve specified stabilization and slewing maneuvers or combinations. The controller includes a thruster-allocation method, which determines the optimal utilization of the available thrusters and introduces redundancy in case of failure. Simulation results are presented for a 3-Unit CubeSat under detumbling, pointing, and pointing and spinning scenarios, as well as comparisons between the thruster-allocation and the paired-firing methods under thruster failure.

Gas discharge ion source. II. Duopigatron

International Nuclear Information System (INIS)

Bacon, F.M.; Bickes, R.W. Jr.; O'Hagan, J.B.

1978-01-01

Ion source performance characteristics consisting of total ion current, ion energy distribution, mass distribution, and ion current density distribution have been measured for several models of a duopigatron. Variations on the duopigatron design involved plasma expansion cup material and dimensions, secondary cathode material, and interelectrode spacings. Of the designs tested, the one with a copper and molybdenum secondary cathode, and a mild steel plasma expansion cup proved to give the best results. The ion current density distribution was peaked at the center of the plasma expansion cup and fell off to 80% of the peak value at the cup wall for a cup 15.2 mm deep. A total ion current of 180 mA consisting of 60%-70% atomic ions was produced with an arc current of 20 A and source pressure of 9.3 Pa. More shallow cups produced a larger beam current and a more sharply peaked ion current density distribution. Typical ion energy distributions were bellshaped curves with a peak 10-20 V below anode potential and with ion energies extending 30-40 V on either side of the peak

Global Linear Stability Analysis of the Spoke Oscillation in Hall Effect Thrusters

Science.gov (United States)

2014-07-15

meνeχ 2 nTe qex (4.1f) ddc dx = 2cpl vix ≡ γ (4.1g) where x is the axial coordinate along the thruster channel; e, me and mi are the electron charge...mi ) P − ( 5 2 Te mi nvex + qex mi ) 1 dc ddc dξ (4.25i) ddc dξ = Pγ (4.25j) Distribution A: Approved for public release; distribution is unlimited...Thruster. PhD thesis, Standford University , 2011. [128] D. Liu, R.E. Huffman, R.D. Branam, and W.A. Hargus. Ultrahigh-speed imaging of hall-thruster

Current Driven Instabilities and Anomalous Mobility in Hall-effect Thrusters

Science.gov (United States)

Tran, Jonathan; Eckhardt, Daniel; Martin, Robert

2017-10-01

Due to the extreme cost of fully resolving the Debye length and plasma frequency, hybrid plasma simulations utilizing kinetic ions and quasi-steady state fluid electrons have long been the principle workhorse methodology for Hall-effect thruster (HET) modeling. Plasma turbulence and the resulting anomalous electron transport in HETs is a promising candidate for developing predictive models for the observed anomalous transport. In this work, we investigate the implementation of an anomalous electron cross field transport model for hybrid HET simulations such a HPHall. A theory for anomalous transport in HETs and current driven instabilities has been recently studied by Lafleur et al. This work has shown collective electron-wave scattering due to large amplitude azimuthal fluctuations of the electric field. We will further adapt the previous results for related current driven instabilities to electric propulsion relevant mass ratios and conduct a preliminary study of resolving this instability with a modified hybrid (fluid electron and kinetic ion) simulation with the hope of integration with established hybrid HET simulations. This work is supported by the Air Force Office of Scientific Research award FA9950-17RQCOR465.

Steady-State Ion Beam Modeling with MICHELLE

Science.gov (United States)

Petillo, John

2003-10-01

There is a need to efficiently model ion beam physics for ion implantation, chemical vapor deposition, and ion thrusters. Common to all is the need for three-dimensional (3D) simulation of volumetric ion sources, ion acceleration, and optics, with the ability to model charge exchange of the ion beam with a background neutral gas. The two pieces of physics stand out as significant are the modeling of the volumetric source and charge exchange. In the MICHELLE code, the method for modeling the plasma sheath in ion sources assumes that the electron distribution function is a Maxwellian function of electrostatic potential over electron temperature. Charge exchange is the process by which a neutral background gas with a "fast" charged particle streaming through exchanges its electron with the charged particle. An efficient method for capturing this is essential, and the model presented is based on semi-empirical collision cross section functions. This appears to be the first steady-state 3D algorithm of its type to contain multiple generations of charge exchange, work with multiple species and multiple charge state beam/source particles simultaneously, take into account the self-consistent space charge effects, and track the subsequent fast neutral particles. The solution used by MICHELLE is to combine finite element analysis with particle-in-cell (PIC) methods. The basic physics model is based on the equilibrium steady-state application of the electrostatic particle-in-cell (PIC) approximation employing a conformal computational mesh. The foundation stems from the same basic model introduced in codes such as EGUN. Here, Poisson's equation is used to self-consistently include the effects of space charge on the fields, and the relativistic Lorentz equation is used to integrate the particle trajectories through those fields. The presentation will consider the complexity of modeling ion thrusters.

Electric Propellant Solid Rocket Motor Thruster Results Enabling Small Satellites

OpenAIRE

Koehler, Frederick; Langhenry, Mark; Summers, Matt; Villarreal, James; Villarreal, Thomas

2017-01-01

Raytheon Missile Systems has developed and tested true on/off/restart solid propellant thrusters which are controlled only by electrical current. This new patented class of energetic rocket propellant is safe, controllable and simple. The range of applications for this game changing technology includes attitude control systems and a safe alternative to higher impulse space satellite thrusters. Described herein are descriptions and performance data for several small electric propellant solid r...

Experimental studies on the surface confined quiescent plasma at INPE

International Nuclear Information System (INIS)

Ferreira, J.L.; Ferreira, J.G.; Sandonato, G.M.; Alves, M.V.; Ludwig, G.O.; Montes, A.

1988-01-01

The quiescent plasma machines used in several experiments at the Associated Plasma Laboratory in INPE are presented. The research activities comprise particle simulation studies on ion acoustic double layers, and studies on the plasma production and loss in surface confined magnetic multidipole thermionic discharges. Recent results from these studies have shown a non-maxwellian plasma formed in most of the discharge conditions. The plasma leakage through the multidipole fields shows an anomalous diffusion process driven by ion acoustic turbulence in the magnetic sheath. The information derived from these studies are being used in the construction and characterization of ion sources for shallow ion implantation in semiconductors, in ion thruster for space propulsion and in the development of powerful ion sources for future use in neutral beam injection systems. (author) [pt

Experimental studies on the surface confined quiescent plasma at INPE

International Nuclear Information System (INIS)

Ferreira, J.L.; Ferreira, J.G.; Sandonato, G.M.; Alves, M.V.; Ludwig, G.O.; Montes, A.

1988-06-01

Quiescent plasma machines are being used in several experiments at the Associated Plasma Laboratory in INPE. The research activities comprises particle simulation studies on ion acoustic double Layers, and studies on the plasma production and loss in surface confined magnetic multidipole thermionic discharges. Recent results from these studies have shown a non-maxwellian plasma formed in most of the discharge conditions. The plasma leakage through the multidipole fields shows an anomalous diffusion process driven by ion acoustic turbulence in the magnetic sheath. The information derived from these studies are being used in the construction and characterization of ion sources for shallow ion implantation in semiconductors, in ion thruster for space propulsion and in the development of powerful ion sources for future use in neutral beam injection systems. (author) [pt

Contamination Study of Micro Pulsed Plasma Thruster

National Research Council Canada - National Science Library

Kesenek, Ceylan

2008-01-01

.... Micro-Pulsed Plasma Thrusters (PPTs) are highly reliable and simple micro propulsion systems that will offer attitude control, station keeping, constellation flying, and drag compensation for such satellites...

Chaotic waves in Hall thruster plasma

International Nuclear Information System (INIS)

Peradzynski, Zbigniew; Barral, S.; Kurzyna, J.; Makowski, K.; Dudeck, M.

2006-01-01

The set of hyperbolic equations of the fluid model describing the acceleration of plasma in a Hall thruster is analyzed. The characteristic feature of the flow is the existence of a trapped characteristic; i.e. there exists a characteristic line, which never intersects the boundary of the flow region in the thruster. To study the propagation of short wave perturbations, the approach of geometrical optics (like WKB) can be applied. This can be done in a linear as well as in a nonlinear version. The nonlinear version describes the waves of small but finite amplitude. As a result of such an approach one obtains so called transport equation, which are governing the wave amplitude. Due to the existence of trapped characteristics this transport equation appears to have chaotic (turbulent) solutions in both, linear and nonlinear versions

Gas discharge ion source. I. Duoplasmatron

International Nuclear Information System (INIS)

Bacon, F.M.

1978-01-01

The effects of the plasma expansion cup on the operation of a duoplasmatron ion source have been investigated by measuring the total ion current and the distributions of the ion energy, mass, and current density. A copper expansion cup did not affect the magnetic field near the anode of the ion source and consequently the ion current density distribution was sharply peaked near the center of the cup. Ion energy distributions were approximately symmetrical about anode potential. The dominant ionic species were D + 3 and D + at low and high arc currents, respectively. Changes in the electrical potential of the copper cup with respect to the anode produced negligible changes in the above data. A mild steel plasma expansion cup caused the magnetic field to diverge and intercept the cup walls, resulting in ion current density distributions that were flatter and more amenable to focusing than the ones with the copper cup. With the steel cup at anode potential, the ion mass distribution was similar to that from the copper cup; however, the ion energy distribution was asymmetrical about the anode potential with a peak about 10-20 V above anode potential. The total ion current from this mode of operation was about one-third the value from the copper cup. If the steel cup assumed floating potential, about 50 V below anode potential, the total current increased to the level observed from the copper cup and the ion energy distribution was similar to that observed with the copper cup but the current density distribution was much flatter than that of the copper cup. The ion mass distribution was 60%-70% atomic ions over the entire arc current range investigated. Based on these data, a modified plasma expansion cup was designed with tapered steel walls lined with a boron nitride insert. The overall performance of the duoplasmatron ion source with this cup was superior to any of the previous three modes of operation

Theory of a wall sheath in a gas-discharge plasma

International Nuclear Information System (INIS)

Dvinin, S.A.; Dovzhenko, V.A.; Kuzovnikov, A.A.

1999-01-01

An integro-differential equation is proposed that generalizes the plasma-sheath (Langmuir-Tonks) equation to include charge exchange between ions and neutrals in a discharge plasma and makes it possible to correctly analyze how the discharge evolves from the regime of collisionless ion motion to the diffusive regime in pure gases with allowance for the space charge in the sheath at the plasma boundary. The integro-differential equation is solved numerically, and the ionization rate is calculated as a function of the ratio between the ion mean free path and the characteristic discharge dimension. The ion energy distribution function in the positive column of a discharge plasma is computed. The parameter range in which the positive column can exist is examined, and the limits of applicability of different discharge models are analyzed depending on the relations between the ion mean free path, Debye length, and discharge dimension

Measurements of the total energy lost per electron-ion pair lost in low-pressure inductive argon, helium, oxygen and nitrogen discharge

International Nuclear Information System (INIS)

Lee, Young-Kwang; Ku, Ju-Hwan; Chung, Chin-Wook

2011-01-01

Experimental measurements of the total energy lost per electron-ion pair lost, ε T , were performed in a low-pressure inductive atomic gases (Ar, He) and molecular gases (O 2 , N 2 ) discharge. The value of ε T was determined from a power balance based on the electropositive global (volume-averaged) model. A floating harmonic method was employed to measure ion fluxes and electron temperatures at the discharge wall. In the pressure range 5-50 mTorr, it was found that the measured ε T ranged from about 70 to 150 V for atomic gases, but from about 180 to 1300 V for molecular gases. This difference between atomic and molecular discharge is caused by additional collisional energy losses of molecular gases. For argon discharge, the stepwise ionization effect on ε T was observed at relatively high pressures. For different gases, the measured ε T was evaluated with respect to the electron temperature, and then compared with the calculation results, which were derived from collisional and kinetic energy loss. The measured ε T and their calculations showed reasonable agreement.

The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics

Science.gov (United States)

Deans, Matthew C.; Reed, Brian D.; Arrington, Lynn A.; Williams, George J.; Kojima, Jun J.; Kinzbach, McKenzie I.; McLean, Christopher H.

2014-01-01

The Green Propellant Infusion Mission (GPIM) is sponsored by NASA's Space Technology Mission Directorate (STMD) Technology Demonstration Mission (TDM) office. The goal of GPIM is to advance the technology readiness level of a green propulsion system, specifically, one using the monopropellant, AF-M315E, by demonstrating ground handling, spacecraft processing, and on-orbit operations. One of the risks identified for GPIM is potential contamination of sensitive spacecraft surfaces from the effluents in the plumes of AF-M315E thrusters. NASA Glenn Research Center (GRC) is conducting activities to characterize the effects of AF-M315E plume impingement and deposition. GRC has established individual plume models of the 22-N and 1-N thrusters that will be used on the GPIM spacecraft. The model simulations will be correlated with plume measurement data from Laboratory and Engineering Model 22-N, AF-M315E thrusters. The thrusters are currently being tested in a small rocket, altitude facility at NASA GRC. A suite of diagnostics, including Raman spectroscopy, Rayleigh spectroscopy, and Schlieren imaging are being used to acquire plume measurements of AF-M315E thrusters. Plume data will include temperature, velocity, relative density, and species concentration. The plume measurement data will be compared to the corresponding simulations of the plume model. The GRC effort will establish a data set of AF-M315E plume measurements and a plume model that can be used for future AF-M315E applications.

An ion source for radiofrequency-pulsed glow discharge time-of-flight mass spectrometry

International Nuclear Information System (INIS)

González Gago, C.; Lobo, L.; Pisonero, J.; Bordel, N.; Pereiro, R.; Sanz-Medel, A.

2012-01-01

A Grimm-type glow discharge (GD) has been designed and constructed as an ion source for pulsed radiofrequency GD spectrometry when coupled to an orthogonal time of flight mass spectrometer. Pulse shapes of argon species and analytes were studied as a function of the discharge conditions using a new in-house ion source (UNIOVI GD) and results have been compared with a previous design (PROTOTYPE GD). Different behavior and shapes of the pulse profiles have been observed for the two sources evaluated, particularly for the plasma gas ionic species detected. In the more analytically relevant region (afterglow), signals for 40 Ar + with this new design were negligible, while maximum intensity was reached earlier in time for 41 (ArH) + than when using the PROTOTYPE GD. Moreover, while maximum 40 Ar + signals measured along the pulse period were similar in both sources, 41 (ArH) + and 80 (Ar 2 ) + signals tend to be noticeable higher using the PROTOTYPE chamber. The UNIOVI GD design was shown to be adequate for sensitive direct analysis of solid samples, offering linear calibration graphs and good crater shapes. Limits of detection (LODs) are in the same order of magnitude for both sources, although the UNIOVI source provides slightly better LODs for those analytes with masses slightly higher than 41 (ArH) + . - Highlights: ► A new RF-pulsed GD ion source (UNIOVI GD) coupled to TOFMS has been characterized. ► Linear calibration graphs and LODs in the low ppm range are achieved. ► Craters with flat bottoms and vertical walls are obtained. ► UNIOVI source can be easily cleaned as it does not require flow tube. ► UNIOVI GD has a simple design and thus its manufacture is easy and cheap.

Micro-cathode Arc Thruster PhoneSat Experiment

Data.gov (United States)

National Aeronautics and Space Administration — The Micro-cathode Arc Thruster Phonesat Experiment  was a joint project between George Washington University and NASA Ames Research Center that successfully...

The peculiarities of spectral manifestations of high-voltage electric discharge in different phase states of ion systems.

Science.gov (United States)

Gafurov, M M; Aliev, A R; Ataev, M B; Rabadanov, K Sh

2013-10-01

The effects of high-voltage pulsed discharge (HVPD activation) on vibrational spectra of ion salt systems have been studied. The peculiarities of spectral display of HVPD in ion melts and aqueous solutions of electrolytes, in ion-conducting phases of crystalline and glassy salt systems have been investigated. After HVPD a salt system is in non-equilibrium activated state. In the activated state of a salt system, the relaxation time of the vibrational excited states of molecular ions is shorter than in the equilibrium state if the vibrational relaxation rate increases with temperature in the system. For those systems for which the relaxation rate decreases at elevated temperatures, the relaxation time of the vibrational excited states of molecular ions is longer than in the equilibrium state. HVPD activation of a salt system can change the configuration of the electron shell of molecular ions. Therefore, the lifetime values of activated state of salt systems are abnormally large. Copyright © 2013 Elsevier B.V. All rights reserved.

Hot-Fire Testing of a 1N AF-M315E Thruster

Science.gov (United States)

Burnside, Christopher G.; Pedersen, Kevin; Pierce, Charles W.

2015-01-01

This hot-fire test continues NASA investigation of green propellant technologies for future missions. To show the potential for green propellants to replace some hydrazine systems in future spacecraft, NASA Marshall Space Flight Center (MSFC) is continuing to embark on hot-fire test campaigns with various green propellant blends. NASA completed a hot-fire test of a 1N AF-M315E monopropellant thruster at the Marshall Space Flight Center in the small altitude test stand located in building 4205. The thruster is a ground test article used for basic performance determination and catalyst studies. The purpose of the hot-fire testing was for performance determination of a 1N size thruster and form a baseline from which to study catalyst performance and life with follow-on testing to be conducted at a later date. The thruster performed as expected. The result of the hot-fire testing are presented in this paper and presentation.

Plasma simulation in space propulsion : the helicon plasma thruster

OpenAIRE

Navarro Cavallé, Jaume

2017-01-01

The Helicon Plasma Thruster (HPT) is an electrodynamic rocket proposed in the early 2000s. It matches an Helicon Plasma Source (HPS), which ionizes the neutral gas and heats up the plasma, with aMagneticNozzle (MN),where the plasma is supersonically accelerated resulting in thrust. Although the core of this thruster inherits the knowledge on Helicon Plasma sources, dated from the seventies, the HPT technology is still not developed and remains below TRL 4. A deep review of the HPT State-of-ar...

Dynamic Electrochemical Impedance Spectroscopy of a Three-Electrode Lithium-Ion Battery during Pulse Charge and Discharge

International Nuclear Information System (INIS)

Huang, Jun; Ge, Hao; Li, Zhe; Zhang, Jianbo

2015-01-01

Highlights: • Dynamic EIS is performed on a three-electrode pouch cell; • Charge transfer resistance during insertion is generally larger than that during deinsertion due to the surface concentration change; • An inductive behavior is revealed at low frequencies due to the violation of stationary condition in DEIS measurement; • Electrochemical models of a single active particle in both time and frequency domain are developed. • The model predicts a positive correlation between the lower frequency limit and the DC current. - Abstract: The dynamic electrochemical impedance spectroscopy (DEIS) of a three-electrode pouch type lithium-ion battery is measured using a series of sine wave perturbations super-imposed on pulse charge and discharge. The DEIS reveals noticeable differences between charge and discharge at frequencies corresponding to the charge transfer reaction. The charge transfer resistance during intercalation is generally found to be larger than that during deintercalation for the battery chemistry in this study. This result is mainly attributed to the decreased Li ion concentration in the electrolyte during intercalation. At low frequencies, an abnormal inductive behavior is also observed. Such abnormality is found to result from the violation of stationary condition, i.e. the state of the battery under pulse charge or discharge deviates significantly from its initial condition for the perturbation of low frequencies. To analytically define the stationary condition, we develop electrochemical models of a single active particle in both time and frequency domain, which describes the transport of lithium ions in both active particle and electrolyte phase and the interfacial charge transfer reactions at their interface. The lower frequency limit is a key parameter to ensure a quasistationary state during the DEIS measurement. An explicit formulation of the stationary condition predicts a positive correlation between the lower frequency limit and

Creation and behavior of radicals and ions in the Acetylene/Argon microwave ECR discharge

Czech Academy of Sciences Publication Activity Database

Pokorný, Petr; Musil, Jindřich; Novotný, Michal; Lančok, Ján; Fitl, Přemysl; Vlček, Jan

2017-01-01

Roč. 14, č. 12 (2017), s. 1-9, č. článku 1700062. ISSN 1612-8850 R&D Projects: GA ČR GA17-13427S; GA MŠk LO1409; GA MŠk LM2015088 Institutional support: RVO:68378271 Keywords : acetylene * ECR discharge * ions * mass spectrometry * radicals Subject RIV: BM - Solid Matter Physics ; Magnetism OBOR OECD: Condensed matter physics (including formerly solid state physics, supercond.) Impact factor: 2.846, year: 2016

Polyatomic ions from a high current ion implanter driven by a liquid metal ion source

Science.gov (United States)

Pilz, W.; Laufer, P.; Tajmar, M.; Böttger, R.; Bischoff, L.

2017-12-01

High current liquid metal ion sources are well known and found their first application as field emission electric propulsion thrusters in space technology. The aim of this work is the adaption of such kind of sources in broad ion beam technology. Surface patterning based on self-organized nano-structures on, e.g., semiconductor materials formed by heavy mono- or polyatomic ion irradiation from liquid metal (alloy) ion sources (LMAISs) is a very promising technique. LMAISs are nearly the only type of sources delivering polyatomic ions from about half of the periodic table elements. To overcome the lack of only very small treated areas by applying a focused ion beam equipped with such sources, the technology taken from space propulsion systems was transferred into a large single-end ion implanter. The main component is an ion beam injector based on high current LMAISs combined with suited ion optics allocating ion currents in the μA range in a nearly parallel beam of a few mm in diameter. Different types of LMAIS (needle, porous emitter, and capillary) are presented and characterized. The ion beam injector design is specified as well as the implementation of this module into a 200 kV high current ion implanter operating at the HZDR Ion Beam Center. Finally, the obtained results of large area surface modification of Ge using polyatomic Bi2+ ions at room temperature from a GaBi capillary LMAIS will be presented and discussed.

Hardware in the Loop Testing of an Iodine-Fed Hall Thruster

Science.gov (United States)

Polzin, Kurt A.; Peeples, Steven R.; Cecil, Jim; Lewis, Brandon L.; Molina Fraticelli, Jose C.; Clark, James P.

2015-01-01

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload,1 providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature ( less than100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum

Predictive fault-tolerant control of an all-thruster satellite in 6-DOF motion via neural network model updating

Science.gov (United States)

Tavakoli, M. M.; Assadian, N.

2018-03-01

The problem of controlling an all-thruster spacecraft in the coupled translational-rotational motion in presence of actuators fault and/or failure is investigated in this paper. The nonlinear model predictive control approach is used because of its ability to predict the future behavior of the system. The fault/failure of the thrusters changes the mapping between the commanded forces to the thrusters and actual force/torque generated by the thruster system. Thus, the basic six degree-of-freedom kinetic equations are separated from this mapping and a set of neural networks are trained off-line to learn the kinetic equations. Then, two neural networks are attached to these trained networks in order to learn the thruster commands to force/torque mappings on-line. Different off-nominal conditions are modeled so that neural networks can detect any failure and fault, including scale factor and misalignment of thrusters. A simple model of the spacecraft relative motion is used in MPC to decrease the computational burden. However, a precise model by the means of orbit propagation including different types of perturbation is utilized to evaluate the usefulness of the proposed approach in actual conditions. The numerical simulation shows that this method can successfully control the all-thruster spacecraft with ON-OFF thrusters in different combinations of thruster fault and/or failure.

2D PIC simulations for an EN discharge with magnetized electrons and unmagnetized ions

Science.gov (United States)

Lieberman, Michael A.; Kawamura, Emi; Lichtenberg, Allan J.

2009-10-01

We conducted 2D particle-in-cell (PIC) simulations for an electronegative (EN) discharge with magnetized electrons and unmagnetized ions, and compared the results to a previously developed 1D (radial) analytical model of an EN plasma with strongly magnetized electrons and weakly magnetized ions [1]. In both cases, there is a static uniform applied magnetic field in the axial direction. The 1D radial model mimics the wall losses of the particles in the axial direction by introducing a bulk loss frequency term νL. A special (desired) solution was found in which only positive and negative ions but no electrons escaped radially. The 2D PIC results show good agreement with the 1D model over a range of parameters and indicate that the analytical form of νL employed in [1] is reasonably accurate. However, for the PIC simulations, there is always a finite flux of electrons to the radial wall which is about 10 to 30% of the negative ion flux.[4pt] [1] G. Leray, P. Chabert, A.J. Lichtenberg and M.A. Lieberman, J. Phys. D, accepted for publication 2009.

Ionization Capabilities of Hydronium Ions and High Electric Fields Produced by Atmospheric Pressure Corona Discharge.

Science.gov (United States)

Sato, Natsuhiko; Sekimoto, Kanako; Takayama, Mitsuo

2016-01-01

Atmospheric pressure corona discharge (APCD) was applied to the ionization of volatile organic compounds. The mass spectra of analytes having aromatic, phenolic, anilinic, basic and aliphatic in nature were obtained by using vapor supply and liquid smear supply methods. The vapor supply method mainly gave protonated analytes [A+H] + caused by proton transfer from hydronium ion H 3 O + , except for benzene, toluene and n -hexane that have lower proton affinity. The use of the liquid smear supply method resulted in the formation of molecular ion A ·+ and/or dehydride analyte [A-H] + , according to the nature of analytes used. The formation of A ·+ without fragment ions could be explained by the electron tunneling via high electric fields 10 8  V/m at the tip of the corona needle. The dehydride analytes [A-H] + observed in the mass spectra of n -hexane, di- and tributylamines may be explained by the hydride abstraction from the alkyl chains by the hydronium ion. The hydronium ion can play the two-roles for analytes, i.e. , the proton donor to form [A+H] + and the hydride acceptor to form [A-H] + .

Interpretation of duoplasmatron-type ion sources from a model of the discharge

International Nuclear Information System (INIS)

Lejeune, C.

1971-06-01

The performance logical improvement of these sources needs a precise knowledge of the emitting ionized medium, on which the whole of the properties is depending. Ion production mechanisms have been studied in the discharge together with their transport towards the extraction hole. The source properties are described, in a new manner, as a function of the discharge modes. The discharge is characterized by the existence of a mode change, related to anode column neutral atom lowering in anode column (arc starvation). The complementarity of probe measurements and the energy spectra analysis of the charges emitted by the anode hole allowed to get the potential axial profile, to discover an electric energetic electron beam extracted from cathodic plasma by the striction shealth potential difference and to determine the electron density radial profile in the anode column. Result analysis allows to get a simple scheme of plasmas and laws controlling them in each of the important modes. The density and potential axial theoretical repartition has been calculated as a function of independent parameters-anode pressure and arc intensity and of three secondary parameters characterizing the energy exchange (electron temperature) and magnetic field topography. The agreement between model predictions and experimental variations for source properties -more specifically gas nature and geometry- allow to give the duoplasmatrons sources similitude rules. The discharge model has allowed to interpret the luminous emission spectra from the anode column. It has been shown theoretically that the peculiar conditions of ionization and excitation allow to use the column as an amplifier medium in the optically field. This plasma has been used successfully as an active medium for an ionic laser in a continuous mode [fr

Remaining dischargeable time prediction for lithium-ion batteries using unscented Kalman filter

Science.gov (United States)

Dong, Guangzhong; Wei, Jingwen; Chen, Zonghai; Sun, Han; Yu, Xiaowei

2017-10-01

To overcome the range anxiety, one of the important strategies is to accurately predict the range or dischargeable time of the battery system. To accurately predict the remaining dischargeable time (RDT) of a battery, a RDT prediction framework based on accurate battery modeling and state estimation is presented in this paper. Firstly, a simplified linearized equivalent-circuit-model is developed to simulate the dynamic characteristics of a battery. Then, an online recursive least-square-algorithm method and unscented-Kalman-filter are employed to estimate the system matrices and SOC at every prediction point. Besides, a discrete wavelet transform technique is employed to capture the statistical information of past dynamics of input currents, which are utilized to predict the future battery currents. Finally, the RDT can be predicted based on the battery model, SOC estimation results and predicted future battery currents. The performance of the proposed methodology has been verified by a lithium-ion battery cell. Experimental results indicate that the proposed method can provide an accurate SOC and parameter estimation and the predicted RDT can solve the range anxiety issues.

Electrospray Thrusters for Attitude Control of a 1-U CubeSat

Science.gov (United States)

Timilsina, Navin

With a rapid increase in the interest in use of nanosatellites in the past decade, finding a precise and low-power-consuming attitude control system for these satellites has been a real challenge. In this thesis, it is intended to design and test an electrospray thruster system that could perform the attitude control of a 1-unit CubeSat. Firstly, an experimental setup is built to calculate the conductivity of different liquids that could be used as propellants for the CubeSat. Secondly, a Time-Of-Flight experiment is performed to find out the thrust and specific impulse given by these liquids and hence selecting the optimum propellant. On the other hand, a colloidal thruster system for a 1-U CubeSat is designed in Solidworks and fabricated using Lathe and CNC Milling Machine. Afterwards, passive propellant feeding is tested in this thruster system. Finally, the electronic circuit and wireless control system necessary to remotely control the CubeSat is designed and the final testing is performed. Among the propellants studied, Ethyl ammonium nitrate (EAN) was selected as the best propellant for the CubeSat. Theoretical design and fabrication of the thruster system was performed successfully and so was the passive propellant feeding test. The satellite was assembled for the final experiment but unfortunately the microcontroller broke down during the first test and no promising results were found out. However, after proving that one thruster works with passive feeding, it could be said that the ACS testing would have worked if we had performed vacuum compatibility tests for other components beforehand.

Micro Cathode Arc Thruster for PhoneSat: Development and Potential Applications

Science.gov (United States)

Gazulla, Oriol Tintore; Perez, Andres Dono; Agasid, Elwood; Uribe, Eddie; Trinh, Greenfield; Keidar, Michael; Teel, George; Haque, Samudra; Lukas, Joseph; Salas, Alberto Guillen;

Optimizing ideal ion propulsion systems depending on the nature of the propellant

Directory of Open Access Journals (Sweden)

Grigore CICAN

2016-12-01

Full Text Available From all accounts the ion thrusters are characterized by the fact that they produce a very high exhaust velocity and specific impulse, sometimes too high for many missions. The exhaust velocity of the ionized particles is a function of the ratio between electrical charge and mass. The obvious solution is the use of ions with low electrical charge – mass ratio, but many of these substances have a corrosive effect on the acceleration grids, they are toxic and hard to store on board the spacecraft. Currently the most used propellant for the ionic propulsion systems is xenon gas having many advantages, but it is expensive when compared to other propellants. The current paper aims to make an optimization study of ideal ion thrusters depending on the nature of the propellant using for studying a significant number of substances. It will study the variation of the performances: force, specific impulse, efficiency, etc for the same power available on board, for the same accelerating voltage and the same ionic current.

Plasma Reactors and Plasma Thrusters Modeling by Ar Complete Global Models

Directory of Open Access Journals (Sweden)

Chloe Berenguer

2012-01-01

Full Text Available A complete global model for argon was developed and adapted to plasma reactor and plasma thruster modeling. It takes into consideration ground level and excited Ar and Ar+ species and the reactor and thruster form factors. The electronic temperature, the species densities, and the ionization percentage, depending mainly on the pressure and the absorbed power, have been obtained and commented for various physical conditions.

Plasma sheath physics and dose uniformity in enhanced glow discharge plasma immersion ion implantation and deposition

International Nuclear Information System (INIS)

Li Liuhe; Li Jianhui; Kwok, Dixon T. K.; Chu, Paul K.; Wang Zhuo

2009-01-01

Based on the multiple-grid particle-in-cell code, an advanced simulation model is established to study the sheath physics and dose uniformity along the sample stage in order to provide the theoretical basis for further improvement of enhanced glow discharge plasma immersion ion implantation and deposition. At t=7.0 μs, the expansion of the sheath in the horizontal direction is hindered by the dielectric cage. The electron focusing effect is demonstrated by this model. Most of the ions at the inside wall of the cage are implanted into the edge of the sample stage and a relatively uniform ion fluence distribution with a large peak is observed at the end. Compared to the results obtained from the previous model, a higher implant fluence and larger area of uniformity are disclosed.

Electrode erosion in arc discharges at atmospheric pressure

Science.gov (United States)

Hardy, T. L.

1985-01-01

An experimental investigation was performed in an effort to measure and increase lifetime of electrodes in an arcjet thruster. The electrode erosion of various anode and cathode materials was measured after tests in an atmospheric pressure nitrogen arc discharge at powers less than 1 kW. A free-burning arc configuration and a constricted arc configuration were used to test the materials. Lanthanum hexboride and thoriated tungsten had low cathode erosion rates while thoriated tungsten and pure tungsten had the lowest anode erosion rates of the materials tested. Anode cooling, reverse gas flow, an external magnetic fields were all found to reduce electrode mass loss.

The microwave thermal thruster and its application to the launch problem

Science.gov (United States)

Parkin, Kevin L. G.

Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and

Synthesis and electrochemical properties of silicon nanosheets by DC arc discharge for lithium-ion batteries.

Science.gov (United States)

Yu, Xiuhong; Xue, Fanghong; Huang, Hao; Liu, Chunjing; Yu, Jieyi; Sun, Yuejun; Dong, Xinglong; Cao, Guozhong; Jung, Youngguan

2014-06-21

Two-dimensional (2D) ultrathin silicon nanosheets (Si NSs) were synthesized by DC arc discharge method and investigated as anode material for Li-ion batteries. The 2D ultrathin characteristics of Si NSs is confirmed by means of transmission electron microscopy (TEM) and atomic force microscopy (AFM). The average size of Si NSs is about 20 nm, with thickness less than 2.5 nm. The characteristic Raman peak of Si NSs is found to have an appreciable (20 nm) shift to low frequency, presumably due to the size effect. The synergistic effects of Ar(+) and H(+) lead to 2D growth of Si NSs under high temperature and energy. Electrochemical analyses reveal that Si NSs anode possesses stable cycling performance and fast diffusion of Li-ions with insertion/extraction processes. Such Si NSs might be a promising candidate for anode of Li-ion batteries.

HiVHAc Thruster Wear and Structural Tests

Data.gov (United States)

National Aeronautics and Space Administration — NASA GRC is developing a 4.5 kW-class Hall propulsion system. This system includes a long life high performance Hall Effect Thruster (HET), a highly efficient...

Ion Exchange Resin and Clay Vitrification by Plasma Discharges

International Nuclear Information System (INIS)

Diaz A, Laura V.; Pacheco S, Joel O.; Pacheco P, Marquidia; Monroy G, Fabiola; Emeterio H, Miguel; Ramos F, Fidel

2006-01-01

The lack of treatment of a low and intermediate level radioactive waste (LILRW) lead us to propose a vitrification process based on a plasma discharge; this technique incorporates LILRW into a matrix glass composed of ceramic clays material. The Mexican Institute of Nuclear Research (ININ), uses an ion exchange resin IRN 150 (styrene-divinilbence copolymer) in the TRIGA MARK III nuclear reactor. The principal objective of this resin is to absorb particles containing heavy metals and low-level radioactive particles. Once the IRN 150 resin filter capacity has been exceeded, it should be replaced and treated as LILRW. In this work, a transferred plasma system was realized to vitrify this resin taking advantage of its high power density, enthalpy and chemical reactivity as well as its rapid quenching and high operation temperatures. In order to characterize the morphological structure of these clay samples, Scanning Electron Microscopy (SEM), X-Ray Diffraction (XRD) and Thermogravimetric analysis (TGA) techniques were applied before and after the plasma treatment

Generation of positive and negative oxygen ions in magnetron discharge during reactive sputtering of alumina

Czech Academy of Sciences Publication Activity Database

Pokorný, Petr; Bulíř, Jiří; Lančok, Ján; Musil, Jindřich; Novotný, Michal

2010-01-01

Roč. 7, č. 11 (2010), s. 910-914 ISSN 1612-8850 R&D Projects: GA AV ČR IAA100100718; GA AV ČR KAN400100653; GA ČR GP202/09/P324 Institutional research plan: CEZ:AV0Z10100522 Keywords : aluminium oxide * ion-energy distribution function * magnetron * mass spectrometry * pulsed discharges Subject RIV: BM - Solid Matter Physics ; Magnetism Impact factor: 1.643, year: 2010

Investigations of electropositive and electronegative RF discharges

International Nuclear Information System (INIS)

Bryant, P.M.

2000-01-01

Electronegative RF discharges are extensively used in the semi-conductor industry for material processing. Despite this the subject of electronegative RF discharges has been largely neglected. The aim of this thesis is to investigate a RF oxygen discharge by mass/energy spectrometry, a retarding field analyser and an actively compensated Langmuir probe. Measurements are also obtained in argon for comparison. In this thesis pure oxygen will be used as this has relatively simple discharge chemistry with most of the rate constants well known. Ion energy analysis (Chapter 3) show the discharge to contract into the centre of the chamber at low pressures in both gases. The expected thinner peak of the oxygen ion energy distribution was not observed, this is shown to be due to RF modulation of the positive ions with collisions playing a role. The dominant positive ion in the discharge bulk and colliding in the sheath in oxygen was found to be O 2 + with less than 10% O + over the range of pressure investigated (Chapter 4). Various minor ions such as O 3 + and O 4 + were also observed. By actively compensating a Langmuir probe for the first three plasma harmonics it is shown that it is unnecessary to compensate when the amplitude of a given harmonic is comparable to the electron temperature (Chapter 5). A study of Langmuir probe measurements in argon (Chapter 7) has shown that the use of the collisionless Alien, Boyd and Reynolds theory leads to discrepancies in the measured electron densities. The correct density can be obtained by using the perturbation method of Shih and Levi, this corrects for ion-neutral collisions in electropositive plasmas only. This theory is extended to electronegative plasmas (Chapter 6) so that measurements of the negative ion density obtained from the collisionless theory of Arnemiya, Annaratone and Alien can be corrected. Langmuir probe measurements in oxygen indicate a peak in the negative ion density at around 3Pa and are found to be in good

High Fidelity Modeling of Field-Reversed Configuration (FRC) Thrusters (Briefing Charts)

Science.gov (United States)

2017-05-24

THRUSTERS (Briefing Charts) Robert Martin , Eder Sousa, Jonathan Tran Air Force Research Laboratory (AFMC) AFRL/RQRS 1 Ara Drive Edwards AFB, CA 93524... Martin N/A HIGH FIDELITY MODELING OF FIELD-REVERSED CONFIGURATION (FRC) THRUSTERS Robert Martin1, Eder Sousa2, Jonathan Tran2 1AIR FORCE RESEARCH...Distribution is unlimited. PA Clearance No. 17314 MARTIN , SOUSA, TRAN (AFRL/RQRS) DISTRIBUTION A - APPROVED FOR PUBLIC RELEASE; DISTRIBUTION UNLIMITED. PA

Mechanical and tribological properties of AISI 304 stainless steel nitrided by glow discharge compared to ion implantation and plasma immersion ion implantation

International Nuclear Information System (INIS)

Foerster, C.E.; Serbena, F.C.; Silva, S.L.R. da; Lepienski, C.M.; Siqueira, C.J. de M.; Ueda, M.

2007-01-01

Results about mechanical and tribological behavior of AISI 304 stainless steel nitrided by three different ion beam processes - glow discharge (GD), ion implantation (II) and plasma immersion ion implantation (PI3) are reported. Expanded austenite γ N and nitrides phases (Fe 2+x N, γ'-Fe 4 N and Cr-N) were identified as a function of nitriding conditions. Hardness (H) and elastic modulus (E) profiles were obtained by instrumented penetration. The hardness reached values as high as 21 GPa by PI3. Tribological behavior was studied by reciprocating sliding tests with a WC (Co) ball at room temperature (RT) in dry condition. Different wear regimes were identified in the friction coefficient profiles. The profile form and the running-in distance are strongly dependent on the nitriding process. Adhesive and abrasive wear components can be inferred from these friction profiles. Hardness and tribological performance, after the nitriding processes, are discussed in terms of surface microstructure

Electrostatic Discharge Test of Multi-Junction Solar Array Coupons After Combined Space Environmental Exposures

Science.gov (United States)

Wright, Kenneth H.; Schneider, Todd; Vaughn, Jason; Hoang, Bao; Funderburk, Victor V.; Wong, Frankie; Gardiner, George

2010-01-01

A set of multi-junction GaAs/Ge solar array test coupons were subjected to a sequence of 5-year increments of combined environmental exposure tests. The test coupons capture an integrated design intended for use in a geosynchronous (GEO) space environment. A key component of this test campaign is conducting electrostatic discharge (ESD) tests in the inverted gradient mode. The protocol of the ESD tests is based on the ISO/CD 11221, the ISO standard for ESD testing on solar array panels. This standard is currently in its final review with expected approval in 2010. The test schematic in the ISO reference has been modified with Space System/Loral designed circuitry to better simulate the on-orbit operational conditions of its solar array design. Part of the modified circuitry is to simulate a solar array panel coverglass flashover discharge. All solar array coupons used in the test campaign consist of 4 cells. The ESD tests are performed at the beginning of life (BOL) and at each 5-year environment exposure point. The environmental exposure sequence consists of UV radiation, electron/proton particle radiation, thermal cycling, and ion thruster plume. This paper discusses the coverglass flashover simulation, ESD test setup, and the importance of the electrical test design in simulating the on-orbit operational conditions. Results from 5th-year testing are compared to the baseline ESD characteristics determined at the BOL condition.

Note: Ion source design for ion trap systems

Science.gov (United States)

Noriega, J. R.; Quevedo, M.; Gnade, B.; Vasselli, J.

2013-06-01

A small plasma (glow discharge) based ion source and circuit are described in this work. The ion source works by producing a high voltage pulsed discharge between two electrodes in a pressure range of 50-100 mTorr. A third mesh electrode is used for ion extraction. The electrodes are small stainless steel screws mounted in a MACOR ionization chamber in a linear arrangement. The electrode arrangement is driven by a circuit, design for low power operation. This design is a proof of concept intended for applications on small cylindrical ion traps.

Sputtering in a glow discharge ion source - pressure dependence: theory and experiment

International Nuclear Information System (INIS)

Mason, R.S.; Pichilingi, Melanie

1994-01-01

A simplified theoretical expression has been developed for a glow discharge to show how the average cathode erosion rate (expressed as the number of atoms per ion of the total bombarding flux) varies with primary sputter yield, pressure, 'diffusion length' and sputtered atom 'stopping' cross section. An inverse pressure dependence is predicted which correlates well with experiment in the 2 and He, tend to converge. It is suggested that this could be due to a change in the mechanism to self-sputtering. Under constant conditions, the erosion rates of different cathode materials still correlate quite well with the differences in their primary sputter yields. (author)

Optimized Magnetic Nozzles for MPD Thrusters, Phase I

Data.gov (United States)

National Aeronautics and Space Administration — Magnetoplasmadynamic (MPD) thrusters can provide the high-specific impulse, high-power propulsion required to enable ambitious human and robotic exploration missions...

Metal negative ion beam extraction from a radio frequency ion source

Energy Technology Data Exchange (ETDEWEB)

Kanda, S.; Yamada, N.; Kasuya, T.; Romero, C. F. P.; Wada, M.

2015-04-08

A metal ion source of magnetron magnetic field geometry has been designed and operated with a Cu hollow target. Radio frequency power at 13.56 MHz is directly supplied to the hollow target to maintain plasma discharge and induce self-bias to the target for sputtering. The extraction of positive and negative Cu ion beams have been tested. The ion beam current ratio of Cu{sup +} to Ar{sup +} has reached up to 140% when Ar was used as the discharge support gas. Cu{sup −} ion beam was observed at 50 W RF discharge power and at a higher Ar gas pressure in the ion source. Improvement of poor RF power matching and suppression of electron current is indispensable for a stable Cu{sup −} ion beam production from the source.

The source of X-rays and high-charged ions based on moderate power vacuum discharge with laser triggering

Directory of Open Access Journals (Sweden)

Alkhimova Mariya A.

2015-06-01

Full Text Available The source of X-ray radiation with the energy of quanta that may vary in the range hν = 1÷12 keV was developed for studies in X-ray interaction with matter and modification of solid surfaces. It was based on a vacuum spark discharge with the laser triggering. It was shown in our experiments that there is a possibility to adjust X-ray radiation spectrum by changing the configuration of the electrode system when the energy stored in the capacitor is varied within the range of 1÷17 J. A comprehensive study of X-ray imaging and quanta energy was carried out. These experiments were carried out for the case of both direct and reverse polarity of the voltage on the electrodes. Additionally, ion composition of plasma created in a laser-triggered vacuum discharge was analyzed. Highly charged ions Zn(+21, Cu(+20 and Fe(+18 were observed.

Redeposition of sputtered material in a glow-discharge lamp measured by means of an ion microprobe mass analyser

International Nuclear Information System (INIS)

Ferreira, N.P.; Bueger, P.A.

1978-01-01

The redeposition of sputtered material on the target in a Grimm-type glow-discharge lamp was studied by means of an ion microprobe mass analyser (IMMA) using 16 O 2 + ions as bombarding species. The target was an aluminium disc with a cylindrical copper insertion, one mm in diameter. The lamp was operated at currents of 50 mA and 100 mA and a voltage of 1200 V. It is estimated that 17% of the copper atoms sputtered are redeposited and may be resputtered. (orig.) [de

Properties and etching rates of negative ions in inductively coupled plasmas and dc discharges produced in Ar/SF6

International Nuclear Information System (INIS)

Draghici, M.; Stamate, E.

2010-01-01

Negative ion production is investigated in a chamber with transversal magnetic filter operated in dc or inductively coupled plasma (ICP) modes in Ar/SF 6 gas mixtures. Plasma parameters are evaluated by mass spectrometry and Langmuir probe for different discharge conditions. The density ratio of negative ion to electron exceeded 300 in dc mode while it was below 100 in the ICP mode. The possibility to apply a large positive bias to an electrode without affecting the plasma potential and the transition from a negative sheath to anodic glow are also investigated. The etching rates by positive and negative ions are evaluated on silicon substrate for different Ar/SF 6 gas ratios. The etching rate by negative ions was with less than 5% smaller than that by positive ions.

Properties and etching rates of negative ions in inductively coupled plasmas and dc discharges produced in Ar/SF6

DEFF Research Database (Denmark)

Draghici, Mihai; Stamate, Eugen

2010-01-01

of negative ion to electron exceeded 300 in dc mode while it was below 100 in the ICP mode. The possibility to apply a large positive bias to an electrode without affecting the plasma potential and the transition from a negative sheath to anodic glow are also investigated. The etching rates by positive...... and negative ions are evaluated on silicon substrate for different Ar/SF6 gas ratios. The etching rate by negative ions was with less than 5% smaller than that by positive ions.......Negative ion production is investigated in a chamber with transversal magnetic filter operated in dc or inductively coupled plasma (ICP) modes in Ar/SF6 gas mixtures. Plasma parameters are evaluated by mass spectrometry and Langmuir probe for different discharge conditions. The density ratio...

Acoustic Resonance Reaction Control Thruster (ARCTIC), Phase I

Data.gov (United States)

National Aeronautics and Space Administration — ORBITEC proposes to develop and demonstrate the innovative Acoustic Resonance Reaction Control Thruster (ARCTIC) to provide rapid and reliable in-space impulse...

Non-Destructive Monitoring of Charge-Discharge Cycles on Lithium Ion Batteries using 7Li Stray-Field Imaging

Science.gov (United States)

Tang, Joel A.; Dugar, Sneha; Zhong, Guiming; Dalal, Naresh S.; Zheng, Jim P.; Yang, Yong; Fu, Riqiang

2013-01-01

Magnetic resonance imaging provides a noninvasive method for in situ monitoring of electrochemical processes involved in charge/discharge cycling of batteries. Determining how the electrochemical processes become irreversible, ultimately resulting in degraded battery performance, will aid in developing new battery materials and designing better batteries. Here we introduce the use of an alternative in situ diagnostic tool to monitor the electrochemical processes. Utilizing a very large field-gradient in the fringe field of a magnet, stray-field-imaging (STRAFI) technique significantly improves the image resolution. These STRAFI images enable the real time monitoring of the electrodes at a micron level. It is demonstrated by two prototype half-cells, graphite∥Li and LiFePO4∥Li, that the high-resolution 7Li STRAFI profiles allow one to visualize in situ Li-ions transfer between the electrodes during charge/discharge cyclings as well as the formation and changes of irreversible microstructures of the Li components, and particularly reveal a non-uniform Li-ion distribution in the graphite. PMID:24005580

Effects of the ion-solid interaction in glow discharge vapour deposition polymerization of pyromellitic dianhydride

International Nuclear Information System (INIS)

Maggioni, G.; Carturan, S.; Rigato, V.; Pieri, U.

2000-01-01

Low energy He ion bombardment of pyromellitic dianhydride monomer used in glow discharge vapour deposition polymerization (GDVDP) of polyimide coatings and its effects on the film deposition process have been studied. The sublimation of the monomer molecules and the simultaneous formation of a damaged, carbon-rich surface layer on the target are discussed from a theoretical point of view based on simulations of the ion-solid interaction. Optical emission and mass spectrometry have been used to analyse the species emitted from the target. In order to study the time evolution of the PMDA target damage, the deposition rate of monomer molecules has been monitored. FT-IR spectroscopy has been used to determine the molecular damaging of the target monomer and deposited films

Fast 2D fluid-analytical simulation of ion energy distributions and electromagnetic effects in multi-frequency capacitive discharges

Science.gov (United States)

Kawamura, E.; Lieberman, M. A.; Graves, D. B.

2014-12-01

A fast 2D axisymmetric fluid-analytical plasma reactor model using the finite elements simulation tool COMSOL is interfaced with a 1D particle-in-cell (PIC) code to study ion energy distributions (IEDs) in multi-frequency capacitive argon discharges. A bulk fluid plasma model, which solves the time-dependent plasma fluid equations for the ion continuity and electron energy balance, is coupled with an analytical sheath model, which solves for the sheath parameters. The time-independent Helmholtz equation is used to solve for the fields and a gas flow model solves for the steady-state pressure, temperature and velocity of the neutrals. The results of the fluid-analytical model are used as inputs to a PIC simulation of the sheath region of the discharge to obtain the IEDs at the target electrode. Each 2D fluid-analytical-PIC simulation on a moderate 2.2 GHz CPU workstation with 8 GB of memory took about 15-20 min. The multi-frequency 2D fluid-analytical model was compared to 1D PIC simulations of a symmetric parallel-plate discharge, showing good agreement. We also conducted fluid-analytical simulations of a multi-frequency argon capacitively coupled plasma (CCP) with a typical asymmetric reactor geometry at 2/60/162 MHz. The low frequency 2 MHz power controlled the sheath width and sheath voltage while the high frequencies controlled the plasma production. A standing wave was observable at the highest frequency of 162 MHz. We noticed that adding 2 MHz power to a 60 MHz discharge or 162 MHz to a dual frequency 2 MHz/60 MHz discharge can enhance the plasma uniformity. We found that multiple frequencies were not only useful for controlling IEDs but also plasma uniformity in CCP reactors.

Hg+ ion density in low-pressure Ar-Hg discharge plasma used for liquid crystal display back-lighting

International Nuclear Information System (INIS)

Goto, Miki; Arai, Toshihiko

1995-01-01

The positive column of a low-pressure Ar-Hg discharge has been applied as a fluorescent light source for illumination. Many studies on the diagnostics and fundamental mechanisms have been carried out on both the classical fluorescent lamp (d=36 mm) and the compact fluorescent lamp (d=12 mm). On the other hand, a lamp of extremely narrow diameter (usually below 6 mm) has been recently developed for liquid crystal display (LCD) back-lighting and its importance is undoubtedly increasing. Some characteristics or mechanisms of the narrow-diameter lamp may be similar to those of the 36 mm one; however the similarity rule does not hold between them due to the contributions from a stepwise ionization process. Therefore, in order to clarify the excitation mechanism in the narrow-diameter lamp quantitatively, various parameters must be measured directly and some analysis must be done. The Hg + ion density and electron density are important parameters for the purpose of clarifying the excitation mechanism quantitatively. In this work, we have measured the Hg + ion density using the modified absorption method, and the electron density using the probe method in the Ar-Hg discharge of the 4 mm bore tube on bath temperature. Moreover, with combining the modified absorption method and the probe method, the Hg 2 + molecular ion density has been determined

Capacitor discharge engineering

CERN Document Server

Früngel, Frank B A

1976-01-01

High Speed Pulse Technology, Volume III: Capacitor Discharge Engineering covers the production and practical application of capacitor dischargers for the generation and utilization of high speed pulsed of energy in different forms. This nine-chapter volume discusses the principles of electric current, voltage, X-rays, gamma rays, heat, beams of electrons, neutrons and ions, magnetic fields, sound, and shock waves in gases and liquids. Considerable chapters consider the applications of capacitor discharges, such as impulse hardening of steel, ultrapulse welding of precision parts, X-ray flash t

Numerical investigation of two interacting parallel thruster-plumes and comparison to experiment

Science.gov (United States)

Grabe, Martin; Holz, André; Ziegenhagen, Stefan; Hannemann, Klaus

2014-12-01

Clusters of orbital thrusters are an attractive option to achieve graduated thrust levels and increased redundancy with available hardware, but the heavily under-expanded plumes of chemical attitude control thrusters placed in close proximity will interact, leading to a local amplification of downstream fluxes and of back-flow onto the spacecraft. The interaction of two similar, parallel, axi-symmetric cold-gas model thrusters has recently been studied in the DLR High-Vacuum Plume Test Facility STG under space-like vacuum conditions, employing a Patterson-type impact pressure probe with slot orifice. We reproduce a selection of these experiments numerically, and emphasise that a comparison of numerical results to the measured data is not straight-forward. The signal of the probe used in the experiments must be interpreted according to the degree of rarefaction and local flow Mach number, and both vary dramatically thoughout the flow-field. We present a procedure to reconstruct the probe signal by post-processing the numerically obtained flow-field data and show that agreement to the experimental results is then improved. Features of the investigated cold-gas thruster plume interaction are discussed on the basis of the numerical results.

Field emission electric propulsion thruster modeling and simulation

Science.gov (United States)

Vanderwyst, Anton Sivaram

Electric propulsion allows space rockets a much greater range of capabilities with mass efficiencies that are 1.3 to 30 times greater than chemical propulsion. Field emission electric propulsion (FEEP) thrusters provide a specific design that possesses extremely high efficiency and small impulse bits. Depending on mass flow rate, these thrusters can emit both ions and droplets. To date, fundamental experimental work has been limited in FEEP. In particular, detailed individual droplet mechanics have yet to be understood. In this thesis, theoretical and computational investigations are conducted to examine the physical characteristics associated with droplet dynamics relevant to FEEP applications. Both asymptotic analysis and numerical simulations, based on a new approach combining level set and boundary element methods, were used to simulate 2D-planar and 2D-axisymmetric probability density functions of the droplets produced for a given geometry and electrode potential. The combined algorithm allows the simulation of electrostatically-driven liquids up to and after detachment. Second order accuracy in space is achieved using a volume of fluid correction. The simulations indicate that in general, (i) lowering surface tension, viscosity, and potential, or (ii) enlarging electrode rings, and needle tips reduce operational mass efficiency. Among these factors, surface tension and electrostatic potential have the largest impact. A probability density function for the mass to charge ratio (MTCR) of detached droplets is computed, with a peak around 4,000 atoms per electron. High impedance surfaces, strong electric fields, and large liquid surface tension result in a lower MTCR ratio, which governs FEEP droplet evolution via the charge on detached droplets and their corresponding acceleration. Due to the slow mass flow along a FEEP needle, viscosity is of less importance in altering the droplet velocities. The width of the needle, the composition of the propellant, the

Reaction Control System Thruster Cracking Consultation: NASA Engineering and Safety Center (NESC) Materials Super Problem Resolution Team (SPRT) Findings

Science.gov (United States)

MacKay, Rebecca A.; Smith, Stephen W.; Shah, Sandeep R.; Piascik, Robert S.

2005-01-01

The shuttle orbiter s reaction control system (RCS) primary thruster serial number 120 was found to contain cracks in the counter bores and relief radius after a chamber repair and rejuvenation was performed in April 2004. Relief radius cracking had been observed in the 1970s and 1980s in seven thrusters prior to flight; however, counter bore cracking had never been seen previously in RCS thrusters. Members of the Materials Super Problem Resolution Team (SPRT) of the NASA Engineering and Safety Center (NESC) conducted a detailed review of the relevant literature and of the documentation from the previous RCS thruster failure analyses. It was concluded that the previous failure analyses lacked sufficient documentation to support the conclusions that stress corrosion cracking or hot-salt cracking was the root cause of the thruster cracking and lacked reliable inspection controls to prevent cracked thrusters from entering the fleet. The NESC team identified and performed new materials characterization and mechanical tests. It was determined that the thruster intergranular cracking was due to hydrogen embrittlement and that the cracking was produced during manufacturing as a result of processing the thrusters with fluoride-containing acids. Testing and characterization demonstrated that appreciable environmental crack propagation does not occur after manufacturing.

Application of MEVVA discharge to material surface modification

International Nuclear Information System (INIS)

Gao Yu; Geng Man; Huang Yuming; Gong Xiaorong; Yu Yijun; Tang Deli; Tie Jun

1996-01-01

The authors describes some characteristics of the MEVVA discharge, the process of generating a cathode-arc plasma and the advantages of the MEVVA discharge compared with the kind of heating-vaporizing-ionizing source. Some practical parameters and the operating process of the MEVVA ion source as well as a plasma source with MEVVA discharge used in a PSII device are presented. Various plasmas having good-quality and high-performance are obtained with MEVVA discharges and have been widely used in sight-line processing and omnibearing ion implantation for material surface modification

Hydrogen discharges operating at atmospheric pressure in a semiconductor gas discharge system