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Sample records for 30-cm ion thruster

  1. An engineering model 30 cm ion thruster

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.; King, H. J.; Schnelker, D. E.

    1973-01-01

    Thruster development at Hughes Research Laboratories and NASA Lewis Research Center has brought the 30-cm mercury bombardment ion thruster to the state of an engineering model. This thruster has been designed to have sufficient internal strength for direct mounting on gimbals, to weigh 7.3 kg, to operate with a corrected overall efficiency of 71%, and to have 10,000 hours lifetime. Subassemblies, such as the ion optical system, isolators, etc., have been upgraded to meet launch qualification standards. This paper presents a summary of the design specifications and performance characteristics which define the interface between the thruster module and the remainder of the propulsion system.

  2. The 30-cm ion thruster power processor

    NASA Technical Reports Server (NTRS)

    Herron, B. G.; Hopper, D. J.

    1978-01-01

    A power processor unit for powering and controlling the 30 cm Mercury Electron-Bombardment Ion Thruster was designed, fabricated, and tested. The unit uses a unique and highly efficient transistor bridge inverter power stage in its implementation. The system operated from a 200 to 400 V dc input power bus, provides 12 independently controllable and closely regulated dc power outputs, and has an overall power conditioning capacity of 3.5 kW. Protective circuitry was incorporated as an integral part of the design to assure failure-free operation during transient and steady-state load faults. The implemented unit demonstrated an electrical efficiency between 91.5 and 91.9 at its nominal rated load over the 200 to 400 V dc input bus range.

  3. Ion accelerator systems for high power 30 cm thruster operation

    NASA Technical Reports Server (NTRS)

    Aston, G.

    1982-01-01

    Two and three-grid accelerator systems for high power ion thruster operation were investigated. Two-grid translation tests show that over compensation of the 30 cm thruster SHAG grid set spacing the 30 cm thruster radial plasma density variation and by incorporating grid compensation only sufficient to maintain grid hole axial alignment, it is shown that beam current gains as large as 50% can be realized. Three-grid translation tests performed with a simulated 30 cm thruster discharge chamber show that substantial beamlet steering can be reliably affected by decelerator grid translation only, at net-to-total voltage ratios as low as 0.05.

  4. Increased capabilities of the 30-cm diameter Hg ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Hawkins, C. E.

    1979-01-01

    Some space flight missions require advanced ion thrusters which operate at conditions much different than those for which the baseline 30-cm Hg thruster was developed. Results of initial tests of a 30-cm Hg thruster with two and three grid ion accelerating systems, operated at higher values of both thrust and power and over a greater range of specific impulse than the baseline conditions are presented. Thruster lifetime at increased input power was evaluated both by extended tests and real time spectroscopic measurements.

  5. Status of 30 cm mercury ion thruster development

    NASA Technical Reports Server (NTRS)

    Sovey, J. S.; King, H. J.

    1974-01-01

    Two engineering model 30-cm ion thrusters were assembled, calibrated, and qualification tested. This paper discusses the thruster design, performance, and power system. Test results include documentation of thrust losses due to doubly charged mercury ions and beam divergence by both direct thrust measurements and beam probes. Diagnostic vibration tests have led to improved designs of the thruster backplate structure, feed system, and harness. Thruster durability is being demonstrated over a thrust range of 97 to 113 mN at a specific impulse of about 2900 seconds. As of August 15, 1974, the thruster has successfully operated for over 4000 hours.

  6. Recycle Requirements for NASA's 30 cm Xenon Ion Thruster

    NASA Technical Reports Server (NTRS)

    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.

  7. Performance and Vibration of 30 cm Pyrolytic Ion Thruster Optics

    NASA Technical Reports Server (NTRS)

    Haag, Thomas; Soulas, George C.

    2004-01-01

    Carbon has a sputter erosion rate about an order of magnitude less than that of molybdenum, over the voltages typically used in ion thruster applications. To explore its design potential, 30 cm pyrolytic carbon ion thruster optics have been fabricated geometrically similar to the molybdenum ion optics used on NSTAR. They were then installed on an NSTAR Engineering Model thruster, and experimentally evaluated over much of the original operating envelope. Ion beam currents ranged from 0.51 to 1.76 Angstroms, at total voltages up to 1280 V. The perveance, electron back-streaming limit, and screen-grid transparency were plotted for these operating points, and compared with previous data obtained with molybdenum. While thruster performance with pyrolytic carbon was quite similar to that with molybdenum, behavior variations can reasonably be explained by slight geometric differences. Following all performance measurements, the pyrolytic carbon ion optics assembly was subjected to an abbreviated vibration test. The thruster endured 9.2 g(sub rms) of random vibration along the thrust axis, similar to DS 1 acceptance levels. Despite significant grid clashing, there was no observable damage to the ion optics assembly.

  8. Compensated control loops for a 30-cm ion thruster

    NASA Technical Reports Server (NTRS)

    Robson, R. R.

    1976-01-01

    The vaporizer dynamic control characteristics of a 30-cm diameter mercury ion thruster were determined by operating the thruster in an open loop steady state mode and then introducing a small sinusoidal signal on the main, cathode, or neutralizer vaporizer current and observing the response of the beam current, discharge voltage, and neutralizer keeper voltage, respectively. This was done over a range of frequencies and operating conditions. From these data, Bode plots for gain and phase were made and mathematical models were obtained. The Bode plots and mathematical models were analyzed for stability and appropriate compensation networks determined. The compensated control loops were incorporated into a power processor and operated with a thruster. The time responses of the compensated loops to changes in set points and recovery from arc conditions are presented.

  9. Direct thrust measurement of a 30-cm ion thruster

    NASA Technical Reports Server (NTRS)

    Banks, B.; Rawlin, V.; Weigand, A. J.; Walker, J.

    1975-01-01

    A direct thrust measurement of a 30-cm diameter ion thruster was accomplished by means of a laser interferometer thrust stand. The thruster was supported in a pendulum manner by three 3.65-m long wires. Electrical power was provided by means of 18 mercury filled pots. A movable 23-button planar probe rake was used to determine thrust loss due to ion beam divergence. Values of thrust, thrust loss due to ion beam divergence, and thrust loss due to multiple ionization were measured for ion beam currents ranging from 0.5 A to 2.5 A. Measured thrust values indicate an accuracy of approximately 1% and are in good agreement with thrust values calculated by indirect measurements.

  10. Retrofit and acceptance test of 30-cm ion thrusters

    NASA Technical Reports Server (NTRS)

    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.

  11. Retrofit and verification test of a 30-cm ion thruster

    NASA Technical Reports Server (NTRS)

    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.

  12. Performance of 30-cm ion thrusters with dished accelerator grids

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1973-01-01

    Thirteen sets of dished accelerator grids were treated on five different 30 cm diameter bombardment thrusters to evaluate the effects of grid geometry variations on thruster discharge chamber performance. The dished grid parameters varied were: grid-to-grid spacing, screen and accelerator grid hole diameter, screen and accelerator open area fraction, compensation for beam divergence losses, and accelerator grid thickness. The effects on discharge chamber performance of main magnetic field changes, magnetic baffle current, cathode pole piece length and cathode position were also investigated.

  13. Digital computer control of a 30-cm mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Low, C. A., Jr.

    1975-01-01

    The major objective was to define the exact role of an onboard spacecraft computer in the control of ion thrusters. An initial computer control system with accurate high speed capability was designed, programmed, and tested with the computer as the sole control element for an operating ion thruster. The command functions and a code format for a spacecraft digital control system were established. A second computer control system was constructed to operate with these functions and format. A throttle program sequence was established and tested. A two thruster array was tested with these computer control systems and the results reported.

  14. Long lifetime hollow cathodes for 30-cm mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Mirtich, M. J.; Kerslake, W. R.

    1976-01-01

    An experimental investigation of hollow cathodes for 30-cm Hg bombardment thrusters was carried out. Both main and neutralizer cathode configurations were tested with both rolled foil inserts coated with low work function material and impregnated porous tungsten inserts. Temperature measurements of an impregnated insert at various positions in the cathode were made. These, along with the cathode thermal profile are presented. A theory for rolled foil and impregnated insert operation and lifetime in hollow cathodes is developed. Several endurance tests, as long as 18000 hours at emission currents of up to 12 amps were attained with no degradation in performance.

  15. Studies of dished accelerator grids for 30-cm ion thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1973-01-01

    Eighteen geometrically different sets of dished accelerator grids were tested on five 30-cm thrusters. The geometric variation of the grids included the grid-to-grid spacing, the screen and accelerator hole diameters and thicknesses, the screen and accelerator open area fractions, ratio of dish depth to dish diameter, compensation, and aperture shape. In general, the data taken over a range of beam currents for each grid set included the minimum total accelerating voltage required to extract a given beam current and the minimum accelerator grid voltage required to prevent electron backstreaming.

  16. Studies of dished accelerator grids for 30-cm ion thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1973-01-01

    Geometrically different sets of dished accelerator grids were tested on five 30-cm thrusters. The geometric variation of the grids included the grid-to-grid spacing, the screen and accelerator hole diameters and thicknesses, the screen and accelerator open area fractions, ratio of dish depth to the dish diameter, compensation, and aperture shape. In general, the data taken over a range of beam currents for each grid set included the minimum total accelerating voltage required to extract a given beam current and the minimum accelerator grid voltage required to prevent electron backstreaming.

  17. Reduced power processor requirements for the 30-cm diameter Hg ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1979-01-01

    An evaluation of simplifications for the thruster power processor interface for a 30 cm Hg ion thruster is presented. Tests on the engine, thruster control, and the power supplies are performed. Reduced power processors requirements are defined and the impact on thruster design, performance, and lifetime are assessed.

  18. Electric prototype power processor for a 30cm ion thruster

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Inouye, L. Y.; Schoenfeld, A. D.

    1977-01-01

    An electrical prototype power processor unit was designed, fabricated and tested with a 30 cm mercury ion engine for primary space propulsion. The power processor unit used the thyristor series resonant inverter as the basic power stage for the high power beam and discharge supplies. A transistorized series resonant inverter processed the remaining power for the low power outputs. The power processor included a digital interface unit to process all input commands and internal telemetry signals so that electric propulsion systems could be operated with a central computer system. The electrical prototype unit included design improvement in the power components such as thyristors, transistors, filters and resonant capacitors, and power transformers and inductors in order to reduce component weight, to minimize losses, and to control the component temperature rise. A design analysis for the electrical prototype is also presented on the component weight, losses, part count and reliability estimate. The electrical prototype was tested in a thermal vacuum environment. Integration tests were performed with a 30 cm ion engine and demonstrated operational compatibility. Electromagnetic interference data was also recorded on the design to provide information for spacecraft integration.

  19. Measurement of sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Weigand, A. J.; Mirtich, M. J.

    1975-01-01

    A study was undertaken to investigate the sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters. Quartz crystal microbalances and fused silica samples were used to analyze the sputtered flux. Spectral transmittance measurements and spectrographic analysis of the samples were made after they were exposed to different thruster effluence by operating the thrusters at various conditions and durations of time. These measurements were used to locate the source of the efflux and determine its accumulated effect at various locations near the thruster. Comparisons of in situ and ex situ transmittance measurements of samples exposed to thruster efflux are also presented.

  20. Evolution and status of the 30-cm engineering model ion thruster

    NASA Technical Reports Server (NTRS)

    Masek, T. D.; Poeschel, R. L.; Collett, C. R.; Schnelker, D. E.

    1976-01-01

    In the past five years the 30-cm ion thruster has developed from infancy to maturity through the joint efforts of the NASA Lewis Research Center (LeRC) and the Hughes Research Laboratories (HRL). The evolution of the 30-cm thruster from the 200-series design to the present 900-series is described. This evolution has included both breadboard and engineering model type thrusters. The evolution description includes functional requirements, design, performance, endurance test results, and major features. The major part of the discussion centers on Hughes-built hardware although NASA LeRC contributions are reflected in the designs.

  1. Thermal analytic model of 30 cm engineering model mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Oglebay, J. C.

    1975-01-01

    A lumped parameter thermal nodal network was developed for a 30 cm Engineering Model Mercury Ion Thruster. The network consists of approximately 100 nodes coded in SINDA format for use on the Univac 1106/1108 computer. This model takes into account internal dissipation, radiation, and conduction as well as environmental heating. A series of tests were performed to simulate a wide range of thermal environments on an operating 30 cm thruster, instrumented to measure the temperature distribution within the thruster. The results of these tests were used to calibrate the analytical model. The analytical model along with comparisons between analytical and experimental results for the various operating conditions are presented.

  2. Low voltage 30-cm ion thruster development. [including performance and structural integrity (vibration) tests

    NASA Technical Reports Server (NTRS)

    King, H. J.

    1974-01-01

    The basic goal was to advance the development status of the 30-cm electron bombardment ion thruster from a laboratory model to a flight-type engineering model (EM) thruster. This advancement included the more conventional aspects of mechanical design and testing for launch loads, weight reduction, fabrication process development, reliability and quality assurance, and interface definition, as well as a relatively significant improvement in thruster total efficiency. The achievement of this goal was demonstrated by the successful completion of a series of performance and structural integrity (vibration) tests. In the course of the program, essentially every part and feature of the original 30-cm Thruster was critically evaluated. These evaluations, led to new or improved designs for the ion optical system, discharge chamber, cathode isolator vaporizer assembly, main isolator vaporizer assembly, neutralizer assembly, packaging for thermal control, electrical terminations and structure.

  3. Sensitivity of 30-cm mercury bombardment ion thruster characteristics to accelerator grid design

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1978-01-01

    The design of ion optics for bombardment thrusters strongly influences overall performance and lifetime. The operation of a 30 cm thruster with accelerator grid open area fractions ranging from 43 to 24 percent, was evaluated and compared with experimental and theoretical results. Ion optics properties measured included the beam current extraction capability, the minimum accelerator grid voltage to prevent backstreaming, ion beamlet diameter as a function of radial position on the grid and accelerator grid hole diameter, and the high energy, high angle ion beam edge location. Discharge chamber properties evaluated were propellant utilization efficiency, minimum discharge power per beam amp, and minimum discharge voltage.

  4. Sputtering phenomena of discharge chamber components in a 30-cm diameter Hg ion thruster

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.; Rawlin, V. K.

    1976-01-01

    Sputtering and deposition rates were measured for discharge chamber components of a 30-cm diameter mercury ion thruster. It was found that sputtering rates of the screen grid and cathode baffle were strongly affected by geometry of the baffle holder. Sputtering rates of the baffle and screen grid were reduced to 80 and 125 A/hr, respectively, by combination of appropriate geometry and materials selections. Sputtering rates such as these are commensurate with thruster lifetimes of 15,000 hours or more. A semiempirical sputtering model showed good agreement with the measured values.

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

    NASA Technical Reports Server (NTRS)

    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.

  6. Design, fabrication, and operation of dished accelerator grids on a 30-cm ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Banks, B. A.; Byers, D. C.

    1972-01-01

    Several closely-space dished accelerator grid systems were fabricated and tested on a 30-cm diameter mercury bombardment thruster and they appear to be a solution to the stringent requirements imposed by the near-term, high-thrust, low specific impulse electric propulsion missions. The grids were simultaneously hydroformed and then simultaneously stress relieved. The ion extraction capability and discharge chamber performance were studied as the total accelerating voltage, the ratio of net-to-total voltage, grid spacing, and dish direction were varied.

  7. Design, fabrication, and operation of dished accelerator grids on a 30-cm ion thruster.

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Banks, B. A.; Byers, D. C.

    1972-01-01

    Several closely-spaced dished accelerator grid systems have been fabricated and tested on a 30-cm diameter mercury bombardment thruster and they appear to be a solution to the stringent requirements imposed by the near-term, high-thrust, low specific impulse electric propulsion missions. The grids were simultaneously hydroformed and then simultaneously stress relieved. The ion extraction capability and discharge chamber performance were studied as the total accelerating voltage, the ratio of net-to-total voltage, grid spacing, and dish direction were varied.

  8. Development Status of the NASA 30-cm Ion Thruster and Power Processor

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Haag, Thomas W.; Hamley, John A.; Mantenieks, Maris A.; Patterson, Michael J.; Pinero, Luis R.; Rawlin, Vincent K.; Kussmaul, Michael T.; Manzella, David H.; Myers, Roger M.

    1994-01-01

    Xenon ion propulsion systems are being developed by NASA Lewis Research Center and the Jet Propulsion Laboratory to provide flight qualification and validation for planetary and earth-orbital missions. In the ground-test element of this program, light-weight (less than 7 kg), 30 cm diameter ion thrusters have been fabricated, and preliminary design verification tests have been conducted. At 2.3 kW, the thrust, specific impulse, and efficiency were 91 mN, 3300 s, and 0.65, respectively. An engineering model thruster is now undergoing a 2000 h wear-test. A breadboard power processor is being developed to operate from an 80 V to 120 V power bus with inverter switching frequencies of 50 kHz. The power processor design is a pathfinder and uses only three power supplies. The projected specific mass of a flight unit is about 5 kg/kW with an efficiency of 0.92 at the full-power of 2.5 kW. Preliminary integration tests of the neutralizer power supply and the ion thruster have been completed. Fabrication and test of the discharge and beam/accelerator power stages are underway.

  9. Performance of Titanium Optics on a NASA 30 Cm Ion Thruster

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Foster, John E.; Patterson, Michael J.

    2000-01-01

    The results of performance tests with two titanium optics sets are presented and compared to those of molybdenum optics. All tests were conducted on a 30 cm ion thruster that was nearly identical to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster design. Optics performance tests were conducted over a thruster input power range of 0.5 to 4.6 kW. Optics performance including impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, near-field beam current density profiles, beam divergence angles, and beam divergence thrust correction factors were determined throughout this power range. The impingement-limited total voltages for titanium optics were within 10 to 55 V of those for molybdenum optics. Electron backstreaming limit magnitude as a function of peak beam current density for both molybdenum and titanium optics were within a few volts of each other, indicating similar hot grid gaps for these two grid materials during steady-state operation. Beam divergence half-angles at 90 percent of the total beam current and thrust correction factors for both titanium optics sets were within one degree and one percent, respectively, of those for molybdenum optics. When thruster power was increased to 2.3 kW immediately following discharge ignition, the titanium screen grid came into contact with the accelerator grid within five minutes of ignition. Relative to molybdenum, titanium's larger thermal expansion and smaller thermal conductivity likely caused the screen grid to thermally expand more relative to the accelerator grid during startup.

  10. Fabrication and verification testing of ETM 30 cm diameter ion thrusters

    NASA Technical Reports Server (NTRS)

    Collett, C.

    1977-01-01

    Engineering model designs and acceptance tests are described for the 800 and 900 series 30 cm electron bombardment thrustors. Modifications to the test console for a 1000 hr verification test were made. The 10,000 hr endurance test of the S/N 701 thruster is described, and post test analysis results are included.

  11. A 30-cm mercury ion thruster performance with a 1 kW capacitor-diode voltage multiplier beam supply

    NASA Technical Reports Server (NTRS)

    Terdan, F. F.; Harrigill, W. T., Jr.

    1978-01-01

    A 1 kW solar array and capacitor-diode voltage multiplier converter (S/A-CDVM) was successfully integrated with a 30 cm diameter mercury ion thruster system to provide ion beam power. Measurements were made to compare steady state and transient response performance of a conventional bridge converter with the S/A-CDVM converter used for the ion beam supply. The ability to recover from screen to accelerator arcs and promptly re-establish stable thruster performance was demonstrated. Solar array transient response to thruster arcing was measured.

  12. Performance Evaluation of Titanium Ion Optics for the NASA 30 cm Ion Thruster

    NASA Technical Reports Server (NTRS)

    Soulas, George C.

    2001-01-01

    The results of performance tests with titanium ion optics were presented and compared to those of molybdenum ion optics. Both titanium and molybdenum ion optics were initially operated until ion optics performance parameters achieved steady state values. Afterwards, performance characterizations were conducted. This permitted proper performance comparisons of titanium and molybdenum ion optics. Ion optics' performance A,as characterized over a broad thruster input power range of 0.5 to 3.0 kW. All performance parameters for titanium ion optics of achieved steady state values after processing 1200 gm of propellant. Molybdenum ion optics exhibited no burn-in. Impingement-limited total voltages for titanium ion optics where up to 55 V greater than those for molybdenum ion optics. Comparisons of electron backstreaming limits as a function of peak beam current density for molybdenum and titanium ion optics demonstrated that titanium ion optics operated with a higher electron backstreaming limit than molybdenum ion optics for a given peak beam current density. Screen grid ion transparencies for titanium ion optics were as much as 3.8 percent lower than those for molybdenum ion optics. Beam divergence half-angles that enclosed 95 percent of the total beam current for titanium ion optics were within 1 to 3 deg. of those for molybdenum ion optics. All beam divergence thrust correction factors for titanium ion optics were within 1 percent of those with molybdenum ion optics.

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

    NASA Technical Reports Server (NTRS)

    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.

  14. Effect of facility background gases on internal erosion of the 30-cm Hg ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Mantenieks, M. A.

    1978-01-01

    Sputtering erosion of the upstream side of the molybdenum screen grid by discharge chamber ions in mercury bombardment thrusters was considered. Data which revealed that the screen grid erosion was very sensitive to the partial pressure of certain background gases in the space simulation vacuum facility were presented along with results of tests conducted to evaluate this effect. It is shown from estimates of the screen grid erosion in space that adequate lifetime for proposed missions exists.

  15. Thermal Characterization of a NASA 30-cm Ion Thruster Operated up to 5 kW

    NASA Technical Reports Server (NTRS)

    SarverVerhey, Timothy R.; Domonkos, Matthew T.; Patterson, Michael J.

    2001-01-01

    A preliminary thermal characterization of a newly-fabricated NSTAR-derived test-bed thruster has recently been performed. The temperature behavior of the rare-earth magnets are reported because of their critical impact on thruster operation. The results obtained to date showed that the magnet temperatures did not exceed the stabilization Emit during thruster operation up to 4.6 kW. Magnet temperature data were also obtained for two earlier NSTAR Engineering Model Thrusters and are discussed in this report. Comparison between these thrusters suggests that the test-bed engine in its present condition is able to operate safely at higher power because of the lower discharge losses over the entire operating power range of this engine. However, because of the 'burn-in' behavior of the NSTAR thruster, magnet temperatures are expected to increase as discharge losses increase with accumulated thruster operation. Consequently, a new engineering solution may be required to achieve 5-kW operation with acceptable margin.

  16. The effects of exposure to LN2 temperatures and 2.5 suns solar radiation on 30-cm ion thruster performance

    NASA Technical Reports Server (NTRS)

    Mirtich, M. J.

    1975-01-01

    An experimental test program was developed to demonstrate all 30 cm Hg-ion bombardment thruster functions over the thermal environment of several proposed missions. A 30 cm thruster with grids dished 1.25 cm and instrumented with 31 thermocouples, was placed in a vacuum tank equipped with minus 196 C walls. Cold storage of a thruster was simulated and temperatures as low as minus 100 C were attained on the thruster. The thruster started successfully from these cold conditions. The thruster operating at both half and full beam power was exposed to 2.5 suns on axis solar simulation. Various thruster thermal configurations, used to simulate multiple thruster operation, were tested at the above conditions. The results of these tests are reported herein.

  17. The effects of exposure to LN2 temperatures and 2.5 suns solar radiation on 30-cm ion thruster performance

    NASA Technical Reports Server (NTRS)

    Mirtich, M. J.

    1975-01-01

    An experimental test program was developed to demonstrate all 30 cm Hg-ion bombardment thruster functions over the thermal environment of several proposed missions. A 30 cm thruster with grids dished 1.25 cm and instrumented with 31 thermocouples, was placed in a vacuum tank equipped with -196 C walls. Cold storage of a thruster was simulated and temperatures as low as -100 C were attained on the thruster. The thruster started successfully from these cold conditions. The thruster operating at both half and full beam power was exposed to 2.5 suns on axis solar simulation. Various thruster thermal configurations, used to simulate multiple thruster operation, were tested at the above conditions. The results of these tests are reported herein.

  18. Endurance testing of a 30-cm Kaufman thruster

    NASA Technical Reports Server (NTRS)

    Collett, C. R.

    1973-01-01

    Results of a program to demonstrate lifetime capability of a 30-cm Kaufman ion thruster with a 6000 hour endurance test are described. Included in the program are (1) thruster fabrication, (2) design and construction of a test console containing a transistorized high frequency power processor, and control circuits which provide unattended automatic operation of the thruster, and (3) modification of a vacuum facility to incorporate a frozen mercury collector and permit unattended operation. Four tests ranging in duration from 100 to 1100 hours have been completed. These tests and the resulting thruster modifications are described. The status of the endurance test is also presented.

  19. A multiple thruster array for 30-cm thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Mantenieks, M. A.

    1975-01-01

    The 3.0-m diameter chamber of the 7.6-m diameter by 21.4-m long vacuum tank at NASA LeRC was modified to permit testing of an array of up to six 30-cm thrusters with a variety of laboratory and thermal vacuum bread-board power systems. A primary objective of the Multiple Thruster Array (MTA) program is to assess the impact of multiple thruster operation on individual thruster and power processor requirements. The areas of thruster startup, steady-state operation, throttling, high voltage recycle, thrust vectoring, and shutdown are of special concern. The results of initial tests are reported.

  20. Performance documentation of the engineering model 30-cm diameter thruster

    NASA Technical Reports Server (NTRS)

    Bechtel, R. T.; Rawlin, V. K.

    1976-01-01

    The results of extensive testing of two 30-cm ion thrusters which are virtually identical to the 900 series Engineering Model Thruster in an ongoing 15,000-hour life test are presented. Performance data for the nominal fullpower (2650 W) operating point; performance sensitivities to discharge voltage, discharge losses, accelerator voltage, and magnetic baffle current; and several power throttling techniques (maximum Isp, maximum thrust/power ratio, and two cases in between are included). Criteria for throttling are specified in terms of the screen power supply envelope, thruster operating limits, and control stability. In addition, reduced requirements for successful high voltage recycles are presented.

  1. Extended performance technology study 30-cm thruster

    NASA Technical Reports Server (NTRS)

    Beattie, J. R.

    1983-01-01

    The extended performance technology study was an investigation of advanced discharge chambers and thruster components that were designed to operate under conditions which result in an increase in the thrust and thrust to power ratio of the state of the art J-series thruster. The high level of performance was achieved by a discharge chamber that employs a ring cusp magnetic confinement arrangement and a three grid ion extraction assembly. It is shown that the ring cusp magnetic field geometry confines the plasma to the volume immediately adjacent to the ion extraction assembly. A high emission current hollow cathode that demonstrated operation at an emission current as high as J sub E = 40 A, and measurements which show the breakdown voltage of individual sections of the J-series propellant flow electrical isolator is about 340 V per section are investigated.

  2. Performance mapping of a 30 cm engineering model thruster

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.; Vahrenkamp, R. P.

    1975-01-01

    A 30 cm thruster representative of the engineering model design has been tested over a wide range of operating parameters to document performance characteristics such as electrical and propellant efficiencies, double ion and beam divergence thrust loss, component equilibrium temperatures, operational stability, etc. Data obtained show that optimum power throttling, in terms of maximum thruster efficiency, is not highly sensitive to parameter selection. Consequently, considerations of stability, discharge chamber erosion, thrust losses, etc. can be made the determining factors for parameter selection in power throttling operations. Options in parameter selection based on these considerations are discussed.

  3. Endurance test of a 30-CM-diameter engineering model ion thruster. Task 12: Investigation of thin-film erosion monitors for ion thrusters

    NASA Technical Reports Server (NTRS)

    Beattie, J. R.

    1983-01-01

    An investigation of short term measurement techniques for predicting the wearout of ion thrusters resulting from sputter erosion damage is described. The previously established laminar thin film techniques to provide high precision erosion rate data. However, the erosion rates obtained using this technique are generally substantially higher than those obtained during long term endurance tests (by virtue of the as deposited nature of the thin films), so that the results must be interpreted in a relative sense. Absolute measurements can be performed using a new masked substrate arrangement which was developed during this study. This new technique provides a means for estimating the lifetimes of critical discharge chamber components based on direct measurements of sputter erosion depths obtained during short duration (10 hour) tests. The method enables the effects on lifetime of thruster design and operating parameters to be inferred without the investment of the time and capital required to conduct long term (1000 hour) endurance tests. Results obtained using the direct measurement technique are shown to agree with sputter erosion depths calculated for the plasma conditions of the test and also with lifetest results. The direct measurement approach is shown to be applicable to both mercury and argon discharge plasma environments and should be useful in estimating the lifetimes of inert gas and extended performance mercury ion thrusters presently under development.

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

    NASA Technical Reports Server (NTRS)

    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.

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

    NASA Technical Reports Server (NTRS)

    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.

  6. Control of a 30 cm diameter mercury bombardment thruster

    NASA Technical Reports Server (NTRS)

    Terdan, F. F.; Bechtel, R. T.

    1973-01-01

    Increased thruster performance has made closed-loop automatic control more difficult than previously. Specifically, high perveance optics tend to make reliable recycling more difficult. Control logic functions were established for three automatic modes of operation of a 30-cm thruster using a power conditioner console with flight-like characteristics. The three modes provide (1) automatic startup to reach thermal stability, (2) steady-state closed-loop control, and (3) the reliable recycling of the high voltages following an arc breakdown to reestablish normal operation. Power supply impedance characteristics necessary for stable operation and the effect of the magnetic baffle on the reliable recycling was studied.

  7. Translation Optics for 30 cm Ion Engine Thrust Vector Control

    NASA Technical Reports Server (NTRS)

    Haag, Thomas

    2002-01-01

    Data were obtained from a 30 cm xenon ion thruster in which the accelerator grid was translated in the radial plane. The thruster was operated at three different throttle power levels, and the accelerator grid was incrementally translated in the X, Y, and azimuthal directions. Plume data was obtained downstream from the thruster using a Faraday probe mounted to a positioning system. Successive probe sweeps revealed variations in the plume direction. Thruster perveance, electron backstreaming limit, accelerator current, and plume deflection angle were taken at each power level, and for each accelerator grid position. Results showed that the thruster plume could easily be deflected up to six degrees without a prohibitive increase in accelerator impingement current. Results were similar in both X and Y direction.

  8. A 30-cm diameter argon ion source

    NASA Technical Reports Server (NTRS)

    Sovey, J. S.

    1976-01-01

    A 30 cm diameter argon ion source was evaluated. Ion source beam currents up to 4a were extracted with ion energies ranging from 0.2 to 1.5 KeV. An ion optics scaling relation was developed for predicting ion beam extraction capability as a function of total extraction voltage, gas type, and screen grid open area. Ignition and emission characteristics of several hollow cathode geometries were assessed for purposes of defining discharge chamber and neutralizer cathodes. Also presented are ion beam profile characteristics which exhibit broad beam capability well suited for ion beam sputtering applications.

  9. The Use of Laser-Induced Fluorescence to Characterize Discharge Cathode Erosion in a 30 cm Ring-Cusp Ion Thruster

    NASA Technical Reports Server (NTRS)

    Sovey, James S. (Technical Monitor); Williams, George J., Jr.

    2004-01-01

    Relative erosion rates and impingement ion production mechanisms have been identified for the discharge cathode of a 30 cm ion engine using laser-induced fluorescence (LIF). Mo and W erosion products as well as neutral and singly ionized xenon were interrogated. The erosion increased with both discharge current and voltage and spatially resolved measurements agreed with observed erosion patters. Ion velocity mapping identified back-flowing ions near the regions of erosion with energies potentially sufficient to generate the level of observed erosion. Ion production regions downstream of the cathode were indicated and were suggested as possible sources of the erosion causing ions.

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

    NASA Technical Reports Server (NTRS)

    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.

  11. A multiple thruster array for 30-cm thrusters. [propulsion system performance

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Matenieks, M. A.

    1975-01-01

    A 3.0 m diameter chamber of the 7.6 m diameter by 21.4 m long vacuum tank was modified to permit testing of an array of up to six 30-cm thrusters with a variety of laboratory and thermal vacuum breadboard power systems. A primary objective of the Multiple Thruster Array (MTA) program is to assess the impact of multiple thruster operation on individual thruster and power processor requirements. The areas of thruster startup, steady-state operation, throttling, high voltage recycle, thrust vectoring, and shutdown are of special concern. The results of initial tests are reported.

  12. Preliminary results of the mission profile life test of a 30 cm Hg bombardment thruster

    NASA Technical Reports Server (NTRS)

    Bechtel, R. T.; James, E. L.

    1979-01-01

    Long term tests were performed on a 30 cm Hg bombardment thruster and a power processing unit to determine lifetime characteristics. The thruster performance data and other operational characteristics taken at various times during the test segment are presented and evaluated with the life limiting mechanisms: discharge chamber erosion, deposition and spalling, external erosion, cathode degradation, and propellant isolator leakage. The control algorithms for thruster start up, steady state operation, throttle, detection and correction of off normal conditions, and shutdown are discussed.

  13. Experimentally Determined Plasma Parameters in a 30 cm Ion Engine

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Goebel, Dan; Fitzgerald, Dennis; Owens, Al; Tynan, George; Dorner, Russ

    2004-01-01

    Single planar Langmuir probes and fiber optic probes are used to concurrently measure the plasma properties and neutral density variation in a 30cm diameter ion engine discharge chamber, from the immediate vicinity of the keeper to the near grid plasma region. The fiber optic probe consists of a collimated optical fiber recessed into a double bore ceramic tube fitted with a stainless steel light-limiting window. The optical fiber probe is used to measure the emission intensity of excited neutral xenon for a small volume of plasma, at various radial and axial locations. The single Langmuir probes, are used to generate current-voltage characteristics at a total of 140 spatial locations inside the discharge chamber. Assuming a maxwellian distribution for the electron population, the Langmuir probe traces provide spatially resolved measurements of plasma potential, electron temperature, and plasma density. Data reduction for the NSTAR TH8 and TH15 throttle points indicates an electron temperature range of 1 to 7.9 eV and an electron density range of 4e10 to le13 cm(sup -3), throughout the discharge chamber, consistent with the results in the literature. Plasma potential estimates, computed from the first derivative of the probe characteristic, indicate potential from 0.5V to 11V above the discharge voltage along the thruster centerline. These values are believed to be excessively high due to the sampling of the primary electron population along the thruster centerline. Relative neutral density profiles are also obtained with a fiber optic probe sampling photon flux from the 823.1 nm excited to ground state transition. Plasma parameter measurements and neutral density profiles will be presented as a function of probe location and engine discharge conditions. A discussion of the measured electron energy distribution function will also be presented, with regards to variation from pure maxwellian. It has been found that there is a distinct primary population found along

  14. Mercury ion thruster technology

    NASA Technical Reports Server (NTRS)

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

    1989-01-01

    The Mercury Ion Thruster Technology program was an investigation for improving the understanding of state-of-the-art mercury ion thrusters. Emphasis was placed on optimizing the performance and simplifying the design of the 30 cm diameter ring-cusp discharge chamber. Thruster performance was improved considerably; the baseline beam-ion production cost of the optimized configuration was reduced to Epsilon (sub i) perspective to 130 eV/ion. At a discharge propellant-utilization efficiency of 95 percent, the beam-ion production cost was reduced to about 155 eV/ion, representing a reduction of about 40 eV/ion over the corresponding value for the 30 cm diameter J-series thruster. Comprehensive Langmuir-probe surveys were obtained and compared with similar measurements for a J-series thruster. A successful volume-averaging scheme was developed to correlate thruster performance with the dominant plasma processes that prevail in the two thruster designs. The average Maxwellian electron temperature in the optimized ring-cusp design is as much as 1 eV higher than it is in the J-series thruster. Advances in ion-extraction electrode fabrication technology were made by improving materials selection criteria, hydroforming and stress-relieving tooling, and fabrications procedures. An ion-extraction performance study was conducted to assess the effect of screen aperture size on ion-optics performance and to verify the effectiveness of a beam-vectoring model for three-grid ion optics. An assessment of the technology readiness of the J-series thruster was completed, and operation of an 8 cm IAPS thruster using a simplified power processor was demonstrated.

  15. HG ion thruster component testing

    NASA Technical Reports Server (NTRS)

    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.

  16. Segmented ion thruster

    NASA Technical Reports Server (NTRS)

    Brophy, John R. (Inventor)

    1993-01-01

    Apparatus and methods for large-area, high-power ion engines comprise dividing a single engine into a combination of smaller discharge chambers (or segments) configured to operate as a single large-area engine. This segmented ion thruster (SIT) approach enables the development of 100-kW class argon ion engines for operation at a specific impulse of 10,000 s. A combination of six 30-cm diameter ion chambers operating as a single engine can process over 100 kW. Such a segmented ion engine can be operated from a single power processor unit.

  17. Krypton ion thruster performance

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Williams, George J., Jr.

    1992-01-01

    Preliminary data were obtained from a 30 cm ion thruster operating on krypton propellant over the input power range of 0.4-5.5 kW. The data are presented, and compared and contrasted to those 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.

  18. Krypton Ion Thruster Performance

    NASA Technical Reports Server (NTRS)

    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.

  19. Discharge Chamber Plasma Structure of a 30-cm NSTAR-Type Ion Engine

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Gallimore, Alec D.

    2006-01-01

    Single Langmuir probe measurements are presented over a two-dimensional array of locations in the near Discharge Cathode Assembly (DCA) region of a 30-cm diameter ring cusp ion thruster over a range of thruster operating conditions encompassing the high-power half of the NASA throttling table. The Langmuir probe data were analyzed with two separate methods. All data were analyzed initially assuming an electron population consisting of Maxwellian electrons only. The on-axis data were then analyzed assuming both Maxwellian and primary electrons. Discharge plasma data taken with beam extraction exhibit a broadening of the higher electron temperature plume boundary compared to similar discharge conditions without beam extraction. The opposite effect is evident with the electron/ion number density as the data without began, extraction appears to be more collimated than the corresponding data with beam extraction. Primary electron energy and number densities are presented for one operating condition giving an order of magnitude of their value and the error associated with this calculation.

  20. Performance Characterization and Vibration Testing of 30-cm Carbon-Carbon Ion Optics

    NASA Technical Reports Server (NTRS)

    Steven Snyder, John; Brophy, John R.

    2004-01-01

    Carbon-based ion optics have the potential to significantly increase the operable life and power ranges of ion thrusters because of reduced erosion rates compared to molybdenum optics. The development of 15-cm and larger diameter grids has encountered many problems, however, not the least of which is the ability to pass vibration testing. JPL has recently developed a new generation of 30-cm carbon-carbon ion optics in order to address these problems and demonstrate the viability of the technology. Perveance, electron backstreaming, and screen grid transparency data are presented for two sets of optics. Vibration testing was successfully performed on two different sets of ion optics with no damage and the results of those tests are compared to models of grid vibrational behavior. It will be shown that the vibration model is a conservative predictor of grid response and can accurately describe test results. There was no change in grid alignment as a result of vibration testing and a slight improvement, if any change at all, in optics performance.

  1. Status of structural analysis of 30 cm diameter ion optics

    NASA Technical Reports Server (NTRS)

    Macrae, Gregory S.; Hering, Gary T.

    1990-01-01

    Three structural finite element programs are compared with theory, experimental data, and each other to evaluate their usefulness for modeling the thermomechanical deflection of ion engine electrodes. Two programs, NASTRAN and MARC, used a Cray XMP and the third, Algor, used an IBM compatible personal computer. The shape of the applied temperature gradient greatly affects off-axis displacement, implying that an accurate temperature distribution is required to analyze new designs. The use of bulk material constants to model the perforated electrodes was investigated. The stress and displacement predictions are shown to be sensitive to the temperature gradient and the Young's modulus, and insensitive to number of nodes, above some minimum value, and the Poisson ratio used. The models are shown to be useful tools for evaluating designs. Experimental measurements of temperatures and displacements was identified as the most critical area.

  2. Derated ion thruster design issues

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Rawlin, Vincent K.

    1991-01-01

    Preliminary activities to develop and refine a lightweight 30 cm engineering model ion thruster are discussed. The approach is to develop a 'derated' ion thruster capable of performing both auxiliary and primary propulsion roles over an input power range of at least 0.5 to 5.0 kilo-W. Design modifications to a baseline thruster to reduce mass and volume are discussed. Performance data over an order of magnitude input power range are presented, with emphasis on the performance impact of engine throttling. Thruster design modifications to optimize performance over specific power envelopes are discussed. Additionally, lifetime estimates based on wear test measurements are made for the operation envelope of the engine.

  3. Near Discharge Cathode Assembly Plasma Potential Measurements in a 30-cm NSTAR Type Ion Engine During Beam Extraction

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Gallimore, Alec D.

    2006-01-01

    Floating emissive probe plasma potential data are presented over a two-dimensional array of locations in the near Discharge Cathode Assembly (DCA) region of a 30-cm diameter ring-cusp ion thruster. Discharge plasma data are presented with beam extraction at throttling conditions comparable to the NASA TH Levels 8, 12, and 15. The operating conditions of the Extended Life Test (ELT) of the Deep Space One (DS1) flight spare ion engine, where anomalous discharge keeper erosion occurred, were TH 8 and TH 12 consequently they are of specific interest in investigating discharge keeper erosion phenomena. The data do not validate the presence of a potential hill plasma structure downstream of the DCA, which has been proposed as a possible erosion mechanism. The data are comparable in magnitude to data taken by other researchers in ring-cusp electron-bombardment ion thrusters. The plasma potential structures are insensitive to thruster throttling level with a minimum as low as 14 V measured at the DCA exit plane and increasing gradually in the axial direction. A sharp increase in plasma potential to the bulk discharge value of 26 to 28 volts, roughly 10 mm radially from DCA centerline, was observed. Plasma potential measurements indicate a low-potential plume structure that is roughly 20 mm in diameter emanating from the discharge cathode that may be attributed to a free-standing plasma double layer.

  4. Charged particle measurements on a 30-CM diameter mercury ion engine thrust beam

    NASA Technical Reports Server (NTRS)

    Sellen, J. M., Jr.; Komatsu, G. K.; Hoffmaster, D. K.; Kemp, R. F.

    1974-01-01

    Measurements of both thrust ions and charge exchange ions were made in the beam of a 30 centimeter diameter electron bombardment mercury ion thruster. A qualitative model is presented which describes magnitudes of charge exchange ion formation and motions of these ions in the weak electric field structure of the neutralized thrust beam plasma. Areas of agreement and discrepancy between observed and modeled charge exchange properties are discussed.

  5. Ion beam thruster shield

    NASA Technical Reports Server (NTRS)

    Power, J. L. (Inventor)

    1976-01-01

    An ion thruster beam shield is provided that comprises a cylindrical housing that extends downstream from the ion thruster and a plurality of annular vanes which are spaced along the length of the housing, and extend inwardly from the interior wall of the housing. The shield intercepts and stops all charge exchange and beam ions, neutral propellant, and sputter products formed due to the interaction of beam and shield emanating from the ion thruster outside of a fixed conical angle from the thruster axis. Further, the shield prevents the sputter products formed during the operation of the engine from escaping the interior volume of the shield.

  6. Low-Isp derated ion thruster operation

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.

    1992-01-01

    The performance and lifetime expectations of 30 cm xenon ion thruster technology at low values of specific impulse were evaluated, with emphasis on 1000-2500 s operation. Power levels of up to 2.0 kW, appropriate for auxiliary and orbit maneuvering propulsion, were processed at thrust-to-power ratios up to 57 mN/kW. These tests were conducted using a derated 30 cm ion thruster with high-perveance design two-grid ion optics with xenon propellent. Lifetime projections were made based on a simple analysis of critical component erosion rates, and it was found that a strong correlation exists with the ratio of the specific impulse-to-input power. Under all operating conditions for which the projected thruster lifetime is less than 10,000 hrs, the life-limiting component of this technology is erosion of the accelerator grid due to charge-exchange ions. The use of alternative grid materials such as carbon is estimated to increase useful thruster lifetimes by as much as an order of magnitude and may enable long-life high thrust-density, sub-2500 s Isp operation. The performance and life of the derated thruster appears similar to that of the Russian SPT-100 thruster in the 1.0-2.0 kW, 1600-2000 s operational envelope.

  7. Developing a scalable inert gas ion thruster

    NASA Technical Reports Server (NTRS)

    James, E.; Ramsey, W.; Steiner, G.

    1982-01-01

    Analytical studies to identify and then design a high performance scalable ion thruster operating with either argon or xenon for use in large space systems are presented. The magnetoelectrostatic containment concept is selected for its efficient ion generation capabilities. The iterative nature of the bounding magnetic fields allows the designer to scale both the diameter and length, so that the thruster can be adapted to spacecraft growth over time. Three different thruster assemblies (conical, hexagonal and hemispherical) are evaluated for a 12 cm diameter thruster and performance mapping of the various thruster configurations shows that conical discharge chambers produce the most efficient discharge operation, achieving argon efficiencies of 50-80% mass utilization at 240-310 eV/ion and xenon efficiencies of 60-97% at 240-280 eV/ion. Preliminary testing of the large 30 cm thruster, using argon propellant, indicates a 35% improvement over the 12 cm thruster in mass utilization efficiency. Since initial performance is found to be better than projected, a larger 50 cm thruster is already in the development stage.

  8. Ion thruster performance model

    NASA Technical Reports Server (NTRS)

    Brophy, J. R.

    1984-01-01

    A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature. The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.

  9. High-Power Ion Thruster Technology

    NASA Technical Reports Server (NTRS)

    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.

  10. Titanium Optics for Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Haag, Thomas W.; Patterson, Michael J.; Rawlin, Vincent K.

    1999-01-01

    Ion thruster total impulse capability is limited, in part, by accelerator grid sputter erosion. A development effort was initiated to identify a material with a lower accelerator grid volumetric sputter erosion rate than molybdenum, but that could utilize the present NSTAR thruster grid design and fabrication techniques to keep development costs low, and perform as well as molybdenum optics. After comparing the sputter erosion rates of several atomic materials to that of molybdenum at accelerator voltages, titanium was found to offer a 45% reduction in volumetric erosion rates. To ensure that screen grid sputter erosion rates are not higher at discharge chamber potentials, titanium and molybdenum sputter erosion rates were measured at these potentials. Preliminary results showed only a slightly higher volumetric erosion rate for titanium, so that screen grid erosion is insignificant. A number of material, thermal, and mechanical properties were also examined to identify any fabrication, launch environment, and thruster operation issues. Several titanium grid sets were successfully fabricated. A titanium grid set was mounted onto an NSTAR 30 cm engineering model ion thruster and tested to determine optics performance. The titanium optics operated successfully over the entire NSTAR power range of 0.5 to 2.3 kW. Differences in impingement-limited perveances and electron backstreaming limits were found to be due to a larger cold gap for the titanium optics. Discharge losses for titanium grids were lower than those for molybdenum, likely due to a slightly larger titanium screen grid open area fraction. Radial distributions of beam current density with titanium optics were very similar to those with molybdenum optics at all power levels. Temporal electron backstreaming limit measurements showed that titanium optics achieved thermal equilibrium faster than molybdenum optics.

  11. Inert gas ion thruster

    NASA Technical Reports Server (NTRS)

    Ramsey, W. D.

    1980-01-01

    Inert gas performance with three types of 12 cm diameter magnetoelectrostatic containment (MESC) ion thrusters was tested. The types tested included: (1) a hemispherical shaped discharge chamber with platinum cobalt magnets; (2) three different lengths of the hemispherical chambers with samarium cobalt magnets; and (3) three lengths of the conical shaped chambers with aluminum nickel cobalt magnets. The best argon performance was produced by a 8.0 cm long conical chamber with alnico magnets. The best xenon high mass utilization performance was obtained with the same 8.0 cm long conical thruster. The hemispherical thruster obtained 75 to 87% mass utilization at 185 to 205 eV/ion of singly charged ion equivalent beam.

  12. High reliability cathode heaters for ion thrusters

    NASA Technical Reports Server (NTRS)

    Mueller, L. A.

    1976-01-01

    A number of space missions have been proposed which will utilize 30-cm mercury bombardment ion thrusters and also will require a large number of thruster restarts. A test program was carried out to determine thermal cycle life of several different cathode heater designs. Plasma/flame sprayed heaters and swaged type heaters were tested. Four of the five plasma/flame sprayed heaters tested failed in a comparatively short time. Four tantalum swaged heaters that were brazed to the tantalum cathode tube were successfully tested and met the goals that were set at the start of the test.

  13. High reliability cathode heaters for ion thrusters

    NASA Technical Reports Server (NTRS)

    Mueller, L. A.

    1976-01-01

    A number of space missions were proposed which utilize 30-cm mercury bombardment ion thrusters and also require a large number of thruster restarts. A test program was carried out to determine thermal cycle life of several different cathode heater designs. Plasma/flame sprayed heaters and swaged type heaters were tested. Four of the five plasma/flame sprayed heaters tested failed in a comparatively short time. Four tantalum swaged heaters that were brazed to the tantalum cathode tube were successfully tested and met the goals that were set at the start of the test.

  14. Performance of 10-kW class xenon ion thrusters

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Rawlin, Vincent K.

    1988-01-01

    Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

  15. Measurement of Doubly Charged Ions in Ion Thruster Plumes

    NASA Technical Reports Server (NTRS)

    Williams, George J., Jr.; Domonkos, Matthew T.; Chavez, Joy M.

    2002-01-01

    The ratio of doubly to singly charged ions was measured in the plumes of a 30 cm and of a 40 cm ion thruster. The measured ratio was correlated with observed erosion rates and thruster operating conditions. The measured and calculated erosion rates paralleled variation in the j(sup ++)/j(sup +) ratio and indicated that the erosion was dominated by Xe III. Simple models of cathode potential surfaces which were developed in support of this work were in agreement with this conclusion and provided a predictive capability of the erosion given the ratio of doubly to singly charged ion currents.

  16. Plasma properties in electron-bombardment ion thrusters

    NASA Technical Reports Server (NTRS)

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

    1987-01-01

    The paper describes a technique for computing volume-averaged plasma properties within electron-bombardment ion thrusters, using spatially varying Langmuir-probe measurements. Average values of the electron densities are defined by integrating the spatially varying Maxwellian and primary electron densities over the ionization volume, and then dividing by the volume. Plasma properties obtained in the 30-cm-diameter J-series and ring-cusp thrusters are analyzed by the volume-averaging technique. The superior performance exhibited by the ring-cusp thruster is correlated with a higher average Maxwellian electron temperature. The ring-cusp thruster maintains the same fraction of primary electrons as does the J-series thruster, but at a much lower ion production cost. The volume-averaged predictions for both thrusters are compared with those of a detailed thruster performance model.

  17. Results of a 1000-hour wear test of 30-cm carbon-carbon ion optics

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Brophy, John R.; Anderson, John R.

    2005-01-01

    This paper discusses the results of a 1000-hour wear test intended to determine the erosion resistance and voltage standoff capability of the optics in a relevant environment, i.e. duration testing on an NSTAR-like thruster.

  18. Low-Power Ion Thruster Development Status

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.

    1999-01-01

    An effort is on-going to examine scaling relationships and design criteria for ion propulsion systems, and to address the need for a light weight, low power, high specific impulse propulsion option for small spacecraft. An element of this activity is the development of a low-power (sub-0.5 kW) ion thruster. This development effort has led to the fabrication and preliminary performance assessment of an 8 cm prototype xenon ion thruster operating over an input power envelope of 0.1-0.3 kW. Efficiencies for the thruster vary from 0.31 at 1750 seconds specific impulse at 0.1 kW, to about 0.48 at 2700 seconds specific impulse and 0.3 kW input power. Discharge losses for the thruster over this power range varied from about 320-380 W/A down to about 220-250 W/A. Ion optics performance compare favorably to that obtained with 30 cm ion optics, when scaled for the difference in beam area. The neutralizer, fabricated using 3 mm hollow cathode technology, operated at keeper currents of about 0.2-0.3 A, at a xenon flow rate of about 0.06 mg/s, over the 0.1-0.3 kW thruster input power envelope.

  19. A 5-kW xenon ion thruster lifetest

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Verhey, Timothy R.

    1990-01-01

    The results of the first life test of a high power ring-cusp ion thruster are presented. A 30-cm laboratory model thruster was operated steady-state at a nominal beam power of 5 kW on xenon propellant for approximately 900 hours. This test was conducted to identify life-timing erosion modifications, and to demonstrate operation using simplified power processing. The results from this test are described including the conclusions derived from extensive post-test analyses of the thruster. Modifications to the thruster and ground support equipment, which were incorporated to solve problems identified by the lifetest, are also described.

  20. Inert gas ion thruster development

    NASA Technical Reports Server (NTRS)

    Ramsey, W. D.

    1980-01-01

    Two 12 cm magneto-electrostatic containment (MESC) ion thrusters were performance mapped with argon and xenon. The first, hexagonal, thruster produced optimized performance of 48.5to 79 percent argon mass utilization efficiencies at discharge energies of 240 to 425 eV/ion, respectively, Xenon mass utilization efficiencies of 78 to 95 percent were observed at discharge energies of 220 to 290 eV/ion with the same optimized hexagonal thruster. Changes to the cathode baffle reduced the discharge anode potential during xenon operation from approximately 40 volts to about 30 volts. Preliminary tests conducted with the second, hemispherical, MESC thruster showed a nonuniform anode magnetic field adversely affected thruster performance. This performance degradation was partially overcome by changes in the boundary anode placement. Conclusions drawn the hemispherical thruster tests gave insights into the plasma processes in the MESC discharge that will aid in the design of future thrusters.

  1. High frequency plasma generators for ion thruster applications

    NASA Technical Reports Server (NTRS)

    Divergilio, W. F.; Goede, H.; Komatsu, G. K.; Christensen, T.

    1981-01-01

    Two concepts for high frequency discharge ion thrusters are described. Both sources are designed for use with 30 cm grid sets and argon propellant and utilize multi-cusp permanent magnet geometries for plasma confinement. The RF induction source is a conventional design representing a synthesis of the RIT and multi-cusp concepts. The preliminary data (without system optimization) indicate a discharge efficiency comparable to that obtained in 30 cm hollow cathode multi-cusp argon thrusters. The electron cyclotron heating source is electrodeless and exhibits plasma characteristics which should lead to greatly reduced discharge chamber and screen sputter rates with the optimization of the magnetic fields, microwave frequency, and feed configuration.

  2. Performance and optimization of a derated ion thruster for auxiliary propulsion

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Foster, John E.

    1991-01-01

    The characteristics and implications of use of a derated ion thruster for north-south stationkeeping (NSSK) propulsion are discussed. A derated thruster is a 30 cm diameter primary propulsion ion thruster operated at highly throttled conditions appropriate to NSSK functions. The performance characteristics of a 30 cm ion thruster are presented, emphasizing throttled operation at low specific impulse and high thrust-to-power ratio. Performance data and component erosion are compared to other NSSK ion thrusters. Operations benefits derived from the performance advantages of the derated approach are examined assuming an INTELSAt 7-type spacecraft. Minimum ground test facility pumping capabilities required to maintain facility enhanced accelerator grid erosion at acceptable levels in a lifetest are quantified as a function of thruster operating condition. Approaches to reducing the derated thruster mass and volume are also discussed.

  3. Advanced ion thruster research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1985-01-01

    A series of experiments conducted on a ring cusp magnetic field ion thruster; in which the anode, cathode and discharge chamber backplate were moved relative to the magnetic cusp; are described. Optimum locations for the anode, cathode and backplate which yield the lowest energy cost per plasma ion and highest extracted ion fraction are identified. The results are discussed in terms of simple physical models. The results of preliminary experiments into the operation of hollow cathodes on nitrogen and xenon over a large pressure range (0.1 to 100 Torr) are presented. They show that the cathode discharge transfers from the cathode insert to the exterior edge of the orifice plate as the interelectrode pressure is increased. Experimental evidence showing that a new ion extractor grid concept can be used to stabilize the plasma sheath at the screen grid is presented. This concept, identified by the term constrained sheath optics, is shown to hold ion beamlet divergence and impingement characteristics to stable values as the beamlet current and the net and total accelerating voltages are changed. The current status of a study of beamlet vectoring induced by displacing the accelerator and/or decelerator grids of a three grid ion extraction system relative to the screen grid is discussed.

  4. Advanced ion thruster research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1984-01-01

    A simple model describing the discharge chamber performance of high strength, cusped magnetic field ion thrusters is developed. The model is formulated in terms of the energy cost of producing ions in the discharge chamber and the fraction of ions produced in the discharge chamber that are extracted to form the ion beam. The accuracy of the model is verified experimentally in a series of tests wherein the discharge voltage, propellant, grid transparency to neutral atoms, beam diameter and discharge chamber wall temperature are varied. The model is exercised to demonstrate what variations in performance might be expected by varying discharge chamber parameters. The results of a study of xenon and argon orificed hollow cathodes are reported. These results suggest that a hollow cathode model developed from research conducted on mercury cathodes can also be applied to xenon and argon. Primary electron mean free paths observed in argon and xenon cathodes that are larger than those found in mercury cathodes are identified as a cause of performance differences between mercury and inert gas cathodes. Data required as inputs to the inert gas cathode model are presented so it can be used as an aid in cathode design.

  5. An ion thruster module for primary propulsion systems.

    NASA Technical Reports Server (NTRS)

    King, H. J.; Poeschel, R. L.

    1972-01-01

    The development of a 30 cm thruster module having the operational characteristics, weight, and structural integrity consistent with flight hardware is described. Elements of the program discussed in this paper are selection of an ion optical system design, development of the discharge chamber and its control, and the results of extensive performance mapping tests. The thruster system operates at 2750 sec specific impulse at 69% over-all efficiency and can be throttled from 2.0 A to 0.16 A beam current with a control system requiring a single electrical input. The 1 kV ion beam is formed by a high perveance, two grid ion optical system.

  6. Development of advanced inert-gas ion thrusters

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.

    1983-01-01

    Inert gas ion thruster technology offers the greatest potential for providing high specific impulse, low thrust, electric propulsion on large, Earth orbital spacecraft. The development of a thruster module that can be operated on xenon or argon propellant to produce 0.2 N of thrust at a specific impulse of 3000 sec with xenon propellant and at 6000 sec with argon propellant is described. The 30 cm diameter, laboratory model thruster is considered to be scalable to produce 0.5 N thrust. A high efficiency ring cusp discharge chamber was used to achieve an overall thruster efficiency of 77% with xenon propellant and 66% with argon propellant. Measurements were performed to identify ion production and loss processes and to define critical design criteria (at least on a preliminary basis).

  7. NSTAR Ion Thruster Plume Impact Assessments

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Pencil, Eric J.; Rawlin, Vincent K.; Kussmaul, Michael; Oden, Katessha

    1995-01-01

    Tests were performed to establish 30-cm ion thruster plume impacts, including plume characterizations via near and farfield ion current measurements, contamination, and sputtering assessments. Current density measurements show that 95% of the beam was enclosed within a 22 deg half-angle and that the thrust vector shifted by less than 0.3 deg during throttling from 2.3 to 0.5 kW. The beam flatness parameter was found to be 0.47, and the ratio of doubly charged to singly charged ion current density decreased from 15% at 2.3 kW to 5% at 0.5 kW. Quartz sample erosion measurements showed that the samples eroded at a rate of between 11 and 13 pm/khr at 25 deg from the thruster axis, and that the rate dropped by a factor of four at 40 deg. Good agreement was obtained between extrapolated current densities and those calculated from tantalum target erosion measurements. Quartz crystal microbalance and witness plate measurements showed that ion beam sputtering of the tank resulted in a facility material backflux rate of -10 A/hr in a large space simulation chamber.

  8. Miniature Bipolar Electrostatic Ion Thruster

    NASA Technical Reports Server (NTRS)

    Hartley, Frank T.

    2006-01-01

    The figure presents a concept of a bipolar miniature electrostatic ion thruster for maneuvering a small spacecraft. The ionization device in the proposed thruster would be a 0.1-micron-thick dielectric membrane with metal electrodes on both sides. Small conical holes would be micromachined through the membrane and electrodes. An electric potential of the order of a volt applied between the membrane electrodes would give rise to an electric field of the order of several mega-volts per meter in the submicron gap between the electrodes. An electric field of this magnitude would be sufficient to ionize all the molecules that enter the holes. In a thruster-based on this concept, one or more propellant gases would be introduced into such a membrane ionizer. Unlike in larger prior ion thrusters, all of the propellant molecules would be ionized. This thruster would be capable of bipolar operation. There would be two accelerator grids - one located forward and one located aft of the membrane ionizer. In one mode of operation, which one could denote the forward mode, positive ions leaving the ionizer on the backside would be accelerated to high momentum by an electric field between the ionizer and an accelerator grid. Electrons leaving the ionizer on the front side would be ejected into free space by a smaller accelerating field. The equality of the ion and electron currents would eliminate the need for an additional electron- or ion-emitting device to keep the spacecraft charge-neutral. In another mode of operation, which could denote the reverse mode, the polarities of the voltages applied to the accelerator grids and to the electrodes of the membrane ionizer would be the reverse of those of the forward mode. The reversal of electric fields would cause the ion and electrons to be ejected in the reverse of their forward mode directions, thereby giving rise to thrust in the direction opposite that of the forward mode.

  9. A mission profile life test facility. [for mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    James, E.; Vetrone, R.; Bechtel, R.

    1978-01-01

    A test facility is being prepared for a 16,000 hour mission profile life test of multiple electric propulsion thrust subsystems. The facility will be capable of simultaneously operating three 2.7 kW, 30 cm mercury ion thrusters and their power processing. The facility will permit conduction of a program of long-term tests to document thruster characteristics as a function of time and operating point to allow prediction of thruster performance for any mission profile. The thruster will be tested in a 7m by 10m vacuum chamber. Each thruster will be installed in a separate lock chamber so that it can be extended into, or extracted from the main chamber without violating the vacuum integrity of the other thruster. The thrusters will exhaust into a 3m by 5m frozen mercury target. The target and an array of cryopanels to collect sputtered target material will be liquid nitrogen chilled. Power processor units will be tested in an adjacent 1.5m by 2m vacuum chamber and will be temperature controlled by simulated heat pipes.

  10. Miniature Electrostatic Ion Thruster With Magnet

    NASA Technical Reports Server (NTRS)

    Hartley, Frank T.

    2006-01-01

    A miniature electrostatic ion thruster is proposed that, with one exception, would be based on the same principles as those of the device described in the previous article, "Miniature Bipolar Electrostatic Ion Thruster". The exceptional feature of this thruster would be that, in addition to using electric fields for linear acceleration of ions and electrons, it would use a magnetic field to rotationally accelerate slow electrons into the ion stream to neutralize the ions.

  11. NEXT Ion Thruster Thermal Model

    NASA Technical Reports Server (NTRS)

    VanNoord, Jonathan L.

    2010-01-01

    As the NEXT ion thruster progresses towards higher technology readiness, it is necessary to develop the tools that will support its implementation into flight programs. An ion thruster thermal model has been developed for the latest prototype model design to aid in predicting thruster temperatures for various missions. This model is comprised of two parts. The first part predicts the heating from the discharge plasma for various throttling points based on a discharge chamber plasma model. This model shows, as expected, that the internal heating is strongly correlated with the discharge power. Typically, the internal plasma heating increases with beam current and decreases slightly with beam voltage. The second is a model based on a finite difference thermal code used to predict the thruster temperatures. Both parts of the model will be described in this paper. This model has been correlated with a thermal development test on the NEXT Prototype Model 1 thruster with most predicted component temperatures within 5 to 10 C of test temperatures. The model indicates that heating, and hence current collection, is not based purely on the footprint of the magnet rings, but follows a 0.1:1:2:1 ratio for the cathode-to-conical-to-cylindrical-to-front magnet rings. This thermal model has also been used to predict the temperatures during the worst case mission profile that is anticipated for the thruster. The model predicts ample thermal margin for all of its components except the external cable harness under the hottest anticipated mission scenario. The external cable harness will be re-rated or replaced to meet the predicted environment.

  12. Advanced electrostatic ion thruster for space propulsion

    NASA Technical Reports Server (NTRS)

    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.

  13. High frequency plasma generator for ion thrusters

    NASA Technical Reports Server (NTRS)

    Goede, H.; Divergilio, W. F.; Fosnight, V. V.; Komatsu, G.

    1984-01-01

    The results of a program to experimentally develop two new types of plasma generators for 30 cm electrostatic argon ion thrusters are presented. The two plasma generating methods selected for this study were by radio frequency induction (RFI), operating at an input power frequency of 1 MHz, and by electron cyclotron heating (ECH) at an operating frequency of 5.0 GHz. Both of these generators utilize multiline cusp permanent magnet configurations for plasma confinement and beam profile optimization. The program goals were to develop a plasma generator possessing the characteristics of high electrical efficiency (low eV/ion) and simplicity of operation while maintaining the reliability and durability of the conventional hollow cathode plasma sources. The RFI plasma generator has achieved minimum discharge losses of 120 eV/ion while the ECH generator has obtained 145 eV/ion, assuming a 90% ion optical transparency of the electrostatic acceleration system. Details of experimental tests with a variety of magnet configurations are presented.

  14. Ion accelerator system mounting design and operating characteristics for a 5 kW 30-cm xenon ion engine

    NASA Technical Reports Server (NTRS)

    Aston, Graeme; Brophy, John R.

    1987-01-01

    Results from a series of experiments to determine the effect of accelerator grid mount geometry on the performance of the J-series ion optics assembly are described. Three mounting schemes, two flexible and one rigid, are compared for their relative ion extraction capability over a range of total accelerating voltages. The largest ion beam current, for the maximum total voltage investigated, is shown to occur using one of the flexible grid mounting geometries. However, at lower total voltages and reduced engine input power levels, the original rigid J-series ion optics accelerator grid mounts result in marginally better grid system performance at the same cold interelectrode gap.

  15. NEXT Ion Thruster Performance Dispersion Analyses

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Patterson, Michael J.

    2008-01-01

    The NEXT ion thruster is a low specific mass, high performance thruster with a nominal throttling range of 0.5 to 7 kW. Numerous engineering model and one prototype model thrusters have been manufactured and tested. Of significant importance to propulsion system performance is thruster-to-thruster performance dispersions. This type of information can provide a bandwidth of expected performance variations both on a thruster and a component level. Knowledge of these dispersions can be used to more conservatively predict thruster service life capability and thruster performance for mission planning, facilitate future thruster performance comparisons, and verify power processor capabilities are compatible with the thruster design. This study compiles the test results of five engineering model thrusters and one flight-like thruster to determine unit-to-unit dispersions in thruster performance. Component level performance dispersion analyses will include discharge chamber voltages, currents, and losses; accelerator currents, electron backstreaming limits, and perveance limits; and neutralizer keeper and coupling voltages and the spot-to-plume mode transition flow rates. Thruster level performance dispersion analyses will include thrust efficiency.

  16. Extended Performance 8-cm Mercury Ion Thruster

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1981-01-01

    A slightly modified 8-cm Hg ion thruster demonstrated significant increase in performance. Thrust was increased by almost a factor of five over that of the baseline thruster. Thruster operation with various three grid ion optics configurations; thruster performance as a function of accelerator grid open area, cathode baffle, and cathode orifice size; and a life test of 614 hours at a beam current of 250 mA (17.5 mN thrust) are discussed. Highest thruster efficiency was obtained with the smallest open area accelerator grid. The benefits in efficiency from the low neutral loss grids were mitigated, however, by the limitation such grids place on attainable ion beam current densities. The thruster components suffered negligible weight losses during a life test, which indicated that operation of the 8-cm thruster at extended levels of thrust and power is possible with no significant loss of lifetime.

  17. Ion Thruster Support and Positioning System

    NASA Technical Reports Server (NTRS)

    Haag, Thomas W. (Inventor)

    1998-01-01

    A system for supporting and selectively positioning an ion thruster relative to a surface of a spacecraft includes three angularly spaced thruster support assemblies. Each thruster support assembly includes a frame which has a rotary actuator mounted thereon. The rotary actuator is connected to an actuator member which is rotatably connected to a thruster attachment member connected to a body of the thruster. A stabilizer member is rotatably mounted to the frame and to the thruster attachment member. The thruster is selectively movable in the pitch and yaw directions responsive to movement of the actuator members by the actuators on the thruster support assemblies. A failure of any one actuator on a thruster support assembly will generally still enable limited thruster positioning capability in two directions. In a retracted position the thruster attachment members are held in nested relation in saddles supported on the frames of the thruster support assemblies. The thruster is securely held in the retracted position during periods of high loading such as during launch of the spacecraft.

  18. Status of 30-centimeter-diameter mercury ion thruster isolator development

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1976-01-01

    Results are presented of several 30 cm diameter mercury ion thruster isolator life tests that show that the onset and exponential increase of leakage current problems observed in earlier thruster operations and isolator tests have been solved. A 10,006 hour life test of a main isolator vaporizer operated with no mercury flow at 320 C and 1500 volts was found to have no onset of leakage current during the test. A cathode-isolator vaporizer operated with a mercury discharge at 340 to 360 C and 1200 volts for 18,000 hours, was found to have a small increase of leakage current with time. A 10,000 hour thruster life test exhibited no increase of leakage current during the life test. Isolators have been developed which will satisfy 30 cm mercury ion thruster mission requirements.

  19. Comparison of thermal analytic model with experimental test results for 30-sentimeter-diameter engineering model mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Oglebay, J. C.

    1977-01-01

    A thermal analytic model for a 30-cm engineering model mercury-ion thruster was developed and calibrated using the experimental test results of tests of a pre-engineering model 30-cm thruster. A series of tests, performed later, simulated a wide range of thermal environments on an operating 30-cm engineering model thruster, which was instrumented to measure the temperature distribution within it. The modified analytic model is described and analytic and experimental results compared for various operating conditions. Based on the comparisons, it is concluded that the analytic model can be used as a preliminary design tool to predict thruster steady-state temperature distributions for stage and mission studies and to define the thermal interface bewteen the thruster and other elements of a spacecraft.

  20. Development of a large inert gas ion thruster

    NASA Technical Reports Server (NTRS)

    Steiner, G.

    1982-01-01

    A 30 cm inert gas electrostatic ion thruster has been developed, exhibiting excellent performance. In the development, the effective anode area was reduced by altering the magnetic field geometry to improve plasma containment, consistent with operational stability. The propellant introduction scheme has the effect of 'folding' the discharge chamber without the increased wall loss penalty associated with a longer chamber. These features contribute to a low discharge cost (eV/ion) versus mass utilization characteristic which remains relatively flat even to high mass utilizations.

  1. High Frequency Plasma Generators for Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Divergilio, W. F.; Goede, H.; Fosnight, V. V.

    1981-01-01

    The results of a one year program to experimentally adapt two new types of high frequency plasma generators to Argon ion thrusters and to analytically study a third high frequency source concept are presented. Conventional 30 cm two grid ion extraction was utilized or proposed for all three sources. The two plasma generating methods selected for experimental study were a radio frequency induction (RFI) source, operating at about 1 MHz, and an electron cyclotron heated (ECH) plasma source operating at about 5 GHz. Both sources utilize multi-linecusp permanent magnet configurations for plasma confinement. The plasma characteristics, plasma loading of the rf antenna, and the rf frequency dependence of source efficiency and antenna circuit efficiency are described for the RFI Multi-cusp source. In a series of tests of this source at Lewis Research Center, minimum discharge losses of 220+/-10 eV/ion were obtained with propellant utilization of .45 at a beam current of 3 amperes. Possible improvement modifications are discussed.

  2. Lifetime Assessment of the NEXT Ion Thruster

    NASA Technical Reports Server (NTRS)

    VanNoord, Jonathan L.

    2010-01-01

    Ion thrusters are low thrust, high specific impulse devices with required operational lifetimes on the order of 10,000 to 100,000 hr. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest throttling point. Currently, a NEXT engineering model ion thruster with prototype model ion optics is undergoing a long duration test to determine wear characteristics and establish propellant throughput capability. The NEXT thruster includes many improvements over previous generations of ion thrusters, but two of its component improvements have a larger effect on thruster lifetime. These include the ion optics with tighter tolerances, a masked region and better gap control, and the discharge cathode keeper material change to graphite. Data from the NEXT 2000 hr wear test, the NEXT long duration test, and further analysis is used to determine the expected lifetime of the NEXT ion thruster. This paper will review the predictions for all of the anticipated failure mechanisms. The mechanisms will include wear of the ion optics and cathode s orifice plate and keeper from the plasma, depletion of low work function material in each cathode s insert, and spalling of material in the discharge chamber leading to arcing. Based on the analysis of the NEXT ion thruster, the first failure mode for operation above a specific impulse of 2000 sec is expected to be the structural failure of the ion optics at 750 kg of propellant throughput, 1.7 times the qualification requirement. An assessment based on mission analyses for operation below a specific impulse of 2000 sec indicates that the NEXT thruster is capable of double the propellant throughput required by these missions.

  3. The 2.3 kW Ion Thruster Wear Test

    NASA Technical Reports Server (NTRS)

    Parkes, James; Rawlin, Vincent K.; Sovey, James S.; Kussmaul, Michael J.; Patterson, Michael J.

    1995-01-01

    A 30-cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for auxiliary and primary propulsion on missions of national interest. Specific efforts include thruster design optimizations, component life testing and validation, and performance characterizations. Under this program, the ion thruster will be brought to engineering model development status. This paper describes the results of a 2.3-kW 2000-hour wear test performed to identify life limiting phenomena, measure the performance and characterize the operation of the thruster, and obtain wear, erosion, and surface contamination data. These data are being using as a data base for proceeding with additional life validation tests, and to provide input to flight thruster design requirements.

  4. Parallel plate radiofrequency ion thruster

    NASA Technical Reports Server (NTRS)

    Nakanishi, S.

    1982-01-01

    An 8-cm-diam. argon ion thruster is described. It is operated by applying 100 to 160 Mhz rf power across a thin plasma volume in a strongly divergent static magnetic field. No cathode or electron emitter is required to sustain a continuous wave plasma discharge over a broad range of propellant gas flow. Preliminary results indicate that a large fraction of the incident power is being reflected by impedance mismatching in the coupling structure. Resonance effects due to plasma thickness, magnetic field strength, and distribution are presented. Typical discharge losses obtained to date are 500 to 600 W per beam ampere at extracted beam currents up to 60 mA.

  5. A 100 Hour Wear Test of the NASA NSTAR Ion Thruster

    NASA Technical Reports Server (NTRS)

    Polk, J. E.; Patterson, M. J.; Brown, J. R.; Rawlin, V. K.; Sovey, J. S.; Myers, R. M.; Blandino, J. J.; Goodfellow, K. D.; Garner, C. E.

    1996-01-01

    In a NASA program to validate 30 cm Xenon ion thruster technology for use in planetary missions a combination of analysis and testing is being used to establish engine reliability. Five long duration tests are planned to identify new failure mechanisms and characterize the parameters which drive known damage accumulation failure modes.

  6. Ion-thruster propellant utilization

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1971-01-01

    The evaluation and understanding of maximum propellant utilization, with mercury used as the propellant are presented. The primary-electron region in the ion chamber of a bombardment thruster is analyzed at maximum utilization. The results of this analysis, as well as experimental data from a range of ion-chamber configurations, show a nearly constant loss rate for unionized propellant at maximum utilization over a wide range of total propellant flow rate. The discharge loss level of 1000 eV/ion was used as a definition of maximum utilization, but the exact level of this definition has no effect on the qualitative results and little effect on the quantitative results. There are obvious design applications for the results of this investigation, but the results are particularly significant whenever efficient throttled operation is required.

  7. Rapid evaluation of ion thruster lifetime using optical emission spectroscopy

    NASA Technical Reports Server (NTRS)

    Rock, B. A.; Mantenieks, M. A.; Parsons, M. L.

    1985-01-01

    A major life-limiting phenomenon of electric thrusters is the sputter erosion of discharge chamber components. Thrusters for space propulsion are required to operate for extended periods of time, usually in excess of 10,000 hr. Lengthy and very costly life-tests in high-vacuum facilities have been required in the past to determine the erosion rates of thruster components. Alternative methods for determining erosion rates which can be performed in relatively short periods of time at considerably lower costs are studied. An attempt to relate optical emission intensity from an ion bombarded surface (screen grid) to the sputtering rate of that surface is made. The model used a kinetic steady-state (KSS) approach, balancing the rates of population and depopulation of ten low-lying excited states of the sputtered molybdenum atom (MoI) with those of the ground state to relate the spectral intensities of the various transitions of the MoI to the population densities. Once this is accomplished, the population density can be related to the sputtering rate of the target. Radiative and collisional modes of excitation and decay are considered. Since actual data has not been published for MoI excitation rate and decay constants, semiempirical equations are used. The calculated sputtering rate and intensity is compared to the measured intensity and sputtering rates of the 8 and 30 cm ion thrusters.

  8. Performance of Large Area Xenon Ion Thrusters for Orbit Transfer Missions

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.

    1989-01-01

    Studies have indicated that xenon ion propulsion systems can enable the use of smaller earth-launch vehicles for satellite placement which results in significant cost savings. These analyses have assumed the availability of advanced, high power ion thrusters-operating at about 10 kW or higher. A program was initiated to explore the viability of operating 50 cm diameter ion thrusters at this power level. Operation with several discharge chamber and ion extraction grid set combinations was demonstrated and data were obtained at power levels to 16 kW. Fifty cm diameter thrusters using state of the art 30 cm diameter grids or advanced technology 50 cm diameter grids allow discharge power and beam current densities commensurate with long life at power levels up to 10 kW. In addition, 50 cm diameter thrusters are shown to have potential for growth in thrust and power levels beyond 10 kW.

  9. Recent work on an RF ion thruster

    NASA Technical Reports Server (NTRS)

    Lee, R. Q.; Nakanishi, S.

    1981-01-01

    An experimental investigation of an rf ion thruster using an immersed coupler in an argon discharge is reported. The conical coil, used to couple rf power into the discharge, is placed inside the discharge vessel. The discharge was self-sustained by 100-150 MHz rf power at low environmental pressures. The ion extraction was accomplished by conventional accelerated grid optics from an unoptimized 8 cm diameter ion thruster.

  10. Ion Thruster Discharge Performance Per Magnetic Field Topography

    NASA Technical Reports Server (NTRS)

    Wirz, Richard E.; Goebel, Dan

    2006-01-01

    DC-ION is a detailed computational model for predicting the plasma characteristics of rain-cusp ion thrusters. The advanced magnetic field meshing algorithm used by DC-ION allows precise treatment of the secondary electron flow. This capability allows self-consistent estimates of plasma potential that improves the overall consistency of the results of the discharge model described in Reference [refJPC05mod1]. Plasma potential estimates allow the model to predict the onset of plasma instabilities, and important shortcoming of the previous model for optimizing the design of discharge chambers. A magnetic field mesh simplifies the plasma flow calculations, for both the ions and the secondary electrons, and significantly reduces numerical diffusion that can occur with meshes not aligned with the magnetic field. Comparing the results of this model to experimental data shows that the behavior of the primary electrons, and the precise manner of their confinement, dictates the fundamental efficiency of ring-cusp. This correlation is evident in simulations of the conventionally sized NSTAR thruster (30 cm diameter) and the miniature MiXI thruster (3 cm diameter).

  11. Eight-cm mercury ion thruster system technology

    NASA Technical Reports Server (NTRS)

    1974-01-01

    The technology status of 8 cm diameter electron bombardment ion thrusters is presented. Much of the technology resulting from the 5 cm diameter thruster has been adapted and improved upon to increase the reliability, durability, and efficiency of the 8 cm thruster. Technology discussed includes: dependence of neutralizer tip erosion upon neutralizer flow rate; impregnated and rolled-foil insert cathode performance and life testing; neutralizer position studies; thruster ion beam profile measurements; high voltage pulse ignition; high utilization ion machined accelerator grids; deposition internal and external to the thruster; thruster vectoring systems; thruster cycling life testing and thruster system weights for typical mission applications.

  12. Ion sources for space thrusters (invited)

    NASA Astrophysics Data System (ADS)

    Grigoryan, V. G.

    1996-03-01

    One of the main tasks of the creation of spacecraft power plants is raising the thrust producing jet velocity. Conventional chemical engines create jet velocities in the range of 3000-4500 m/s. This situation can be drastically changed if beams of charged particles accelerated by electric and magnetic fields are used to produce thrust. In such cases practically any jet velocity might be created, which considerably enlarges the number of tasks being fulfilled by spacecraft having such types of thruster. Several types of electric propulsion thrusters exist nowadays. They differ in the principles of acceleration of charged particles, for example, arc jets, magnetic plasma dynamic thrusters, stationary plasma thrusters, pulse thrusters, and ion thrusters. Electric propulsion thrusters are practically the accelerators of charged particles which operate under rather strict requirements concerning energy consumption and lifetime. Since the mid-fifties in Russia there have been intensive studies of practically all types of electric propulsion thrusters, including their tests in space, and beginning with the mid-seventies they have been practically used aboard spacecraft with a long, active lifetime. The study of the physical process involved together with the research design allowed Russian scientists to develop electric propulsion thrusters in the power range from hundreds of watts to tens of kilowatts, with jet velocities between 20000 and 50000 m/s and lifetime more than several thousand hours.

  13. Analytical Ion Thruster Discharge Performance Model

    NASA Technical Reports Server (NTRS)

    Goebel, Dan M.; Wirz, Richard E.; Katz, Ira

    2006-01-01

    A particle and energy balance model of the plasma discharge in magnetic ring-cusp ion thrusters has been developed. The model follows the original work of Brophy in the development of global 0-D discharge models that utilize conservation of particles into and out of the thruster and conservation of energy into the discharge and out of the plasma in the form of charged particles to the walls and beam and plasma radiation. The present model is significantly expanded over Brophy's original work by including self-consistent calculations of the internal neutral pressure, electron temperature, primary electron density, electrostatic ion confinement (due to the ring-cusp fields), plasma potential, discharge stability, and time dependent behavior during recycling. The model only requires information on the thruster geometry, ion optics performance and electrical inputs such as discharge voltage and currents, etc. to produce accurate performance curves of discharge loss versus mass utilization efficiency. The model has been benchmarked against the NEXIS Laboratory Model (LM) and Development Model (DM) thrusters, and successfully predicts the thruster discharge loss as a function of mass utilization efficiency for a variety of thrusters. The discharge performance model will be presented and results showing ion thruster performance and stability given.

  14. Hg ion thruster component testing

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1979-01-01

    Electron bombardment thrusters, under development to provide both auxiliary and primary propulsion functions for a large variety of space missions are tested. Thruster design verification which requires life tests of durations of the order of the time anticipated in space applications, are discussed. The life time and reliability of an electron bombardment thruster is dependent upon the performance of several critical components including cathodes, vaporizers, and isolators. The performances of the cathode, vaporizer, and propellant isolaters during fatigue analyses are examined.

  15. Scaling of Ion Thrusters to Low Power

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Grisnik, Stanley P.; Soulas, George C.

    1998-01-01

    Analyses were conducted to examine ion thruster scaling relationships in detail to determine performance limits, and lifetime expectations for thruster input power levels below 0.5 kW. This was motivated by mission analyses indicating the potential advantages of high performance, high specific impulse systems for small spacecraft. The design and development status of a 0.1-0.3 kW prototype small thruster and its components are discussed. Performance goals include thruster efficiencies on the order of 40% to 54% over a specific impulse range of 2000 to 3000 seconds, with a lifetime in excess of 8000 hours at full power. Thruster technologies required to achieve the performance and lifetime targets are identified.

  16. Mercury ion thruster research, 1977. [plasma acceleration

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1977-01-01

    The measured ion beam divergence characteristics of two and three-grid, multiaperture accelerator systems are presented. The effects of perveance, geometry, net-to-total accelerating voltage, discharge voltage and propellant are examined. The applicability of a model describing doubly-charged ion densities in mercury thrusters is demonstrated for an 8-cm diameter thruster. The results of detailed Langmuir probing of the interior of an operating cathode are given and used to determine the ionization fraction as a function of position upstream of the cathode orifice. A mathematical model of discharge chamber electron diffusion and collection processes is presented along with scaling laws useful in estimating performance of large diameter and/or high specific impluse thrusters. A model describing the production of ionized molecular nitrogen in ion thrusters is included.

  17. Improvement of ion thruster design

    NASA Technical Reports Server (NTRS)

    Carpenter, R. T.

    1986-01-01

    Two types of measurements were performed on ion thrustors equipped with SmCo magnets in either ring cusp or line cusp arrangements. Langmuir probes were used to measure plasma potential, electron density, and electron temperture in all regions inside the thruster. Loss fluxes to various surfaces were determined by measuring the currents to foils attached to or imbedded in the surface. Data were obtained for several sets of discharge voltages and currents. The loss currents were determined from current vs voltage characteristics observed on a transistor curve tracer oscilloscope. Both ion and electron currents were measured to all parts of the walls and to all parts of the cathode assembly using collecting plates. These measurement were also made for various parameter sets. In line cusp configuration the plasma density is essentially as predicted by existing calculations. In the ring cusp arrangement the interior of the plasma contains an inhomogeneous and relatively large magnetic field so the geometry is decidely two-dimensional and the models of Self (1967) and of Kino and Sham (1966) do not agree.

  18. Optical properties of mercury ion thruster exhausts and implications for science instruments

    NASA Technical Reports Server (NTRS)

    Monahan, K. M.; Goldstein, R.

    1974-01-01

    Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. Most of the emitted radiation came from UV lines of excited mercury atoms and ions, with no observable continuum emission. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter, was considered incompatible with simultaneous thruster operation.

  19. Grid Gap Measurement for an NSTAR Ion Thruster

    NASA Technical Reports Server (NTRS)

    Diaz, Esther M.; Soulas, George C.

    2006-01-01

    The change in gap between the screen and accelerator grids of an engineering model NSTAR ion optics assembly was measured during thruster operation with beam extraction. The molybdenum ion optics assembly was mounted onto an engineering model NSTAR ion thruster. The measurement technique consisted of measuring the difference in height of an alumina pin relative to the downstream accelerator grid surface. The alumina pin was mechanically attached to the center aperture of the screen grid and protruded through the center aperture of the accelerator grid. The change in pin height was monitored using a long distance microscope coupled to a digital imaging system. Transient and steady-state hot grid gaps were measured at three power levels: 0.5, 1.5 and 2.3 kW. Also, the change in grid gap was measured during the transition between power levels, and during the startup with high voltage applied just prior to discharge ignition. Performance measurements, such as perveance, electron backstreaming limit and screen grid ion transparency, were also made to confirm that this ion optics assembly performed similarly to past testing. Results are compared to a prior test of 30 cm titanium ion optics.

  20. Physical phenomena in mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1979-01-01

    Experimental tests results demonstrating that reductions in screen grid thickness enhance the performance of ion thruster grids are presented. Shaping of the screen hole cross section is shown on the other hand not to affect performance substantially. The effect of the magnetic field in the vicinity of the hollow cathode on cathode performance is studied and test results are presented that show reductions in keeper voltages of a few volts can be realized by judicious applications of fields on the order of 100 gauss. The plasma downstream of a SERT 2 thruster operating without high voltage is studied. A model describing electron escape from the thruster under these conditions is discussed. A model defining the performance of the baffle aperture of an ion thruster is refined and experimental verification of the model is undertaken.

  1. 3-D Simulations of NSTAR Ion Thruster Plasma Interactions

    NASA Technical Reports Server (NTRS)

    Wang, J.; Brophy, J.; Polk, J.; Brinza, D.

    1996-01-01

    Described is a Particle-in-Cell with Monte Carlo Collision code developed to perform detailed three-dimensional ion thruster simulations. To capture the full kinetic behavior of ion thruster plumes, both the electrons and ions are treated as test particles. Simulation results are given of the NSTAR ion thruster under ground test and in space conditions. Numerical results are compared.

  2. The 15 cm diameter ion thruster research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1974-01-01

    The startup reliability of a 15 cm diameter mercury bombardment ion thruster which employs a pulsed high voltage tickler electrode on the main and neutralizer cathodes is examined. Startup of the thruster is achieved 100% of the time on the main cathode and 98.7% of the time on the neutralizer cathode over a 3640 cycle test. The thruster was started from a 20 C initial condition and operated for an hour at a 600 mA beam current. An energy efficiency of 75% and a propellant utilization efficiency of 77% was achieved over the complete cycle. The effect of a single cusp magnetic field thruster length on its performance is discussed. Guidelines are formulated for the shaping of magnetic field lines in thrusters. A model describing double ion production in mercury discharges is presented. The production route is shown to occur through the single ionic ground state. Photographs of the interior of an operating-hollow cathode are presented. A cathode spot is shown to be present if the cathode is free of low work-function surfaces. The spot is observed if a low work-function oxide coating is applied to the cathode insert. Results show that low work-function oxide coatings tend to migrate during thruster operation.

  3. Mercury ion thruster research, 1978

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1978-01-01

    The effects of 8 cm thruster main and neutralizer cathode operating conditions on cathode orifice plate temperatures were studied. The effects of cathode operating conditions on insert temperature profiles and keeper voltages are presented for three different types of inserts. The bulk of the emission current is generally observed to come from the downstream end of the insert rather than from the cathode orifice plate. Results of a test in which the screen grid plasma sheath of a thruster was probed as the beam current was varied are shown. Grid performance obtained with a grid machined from glass ceramic is discussed. The effects of copper and nitrogen impurities on the sputtering rates of thruster materials are measured experimentally and a model describing the rate of nitrogen chemisorption on materials in either the beam or the discharge chamber is presented. The results of optimization of a radial field thruster design are presented. Performance of this device is shown to be comparable to that of a divergent field thruster and efficient operation with the screen grid biased to floating potential, where its susceptibility to sputter erosion damage is reduced, is demonstrated.

  4. Heaterless ignition of inert gas ion thruster hollow cathodes

    NASA Technical Reports Server (NTRS)

    Schatz, M. F.

    1985-01-01

    Heaterless inert gas ion thruster hollow cathodes were investigated with the aim of reducing ion thruster complexity and increasing ion thruster reliability. Cathodes heated by glow discharges are evaluated for power requirements, flowrate requirements, and life limiting mechanisms. An accelerated cyclic life test is presented.

  5. High Power ECR Ion Thruster Discharge Characterization

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Kamhawi, Hani; Haag, Thomas; Carpenter, Christian; Williams, George W.

    2006-01-01

    Electron cyclotron resonance (ECR) based ion thrusters with carbon based ion optics can potentially satisfy lifetime requirements for long duration missions (approximately 10 years) because grid erosion and cathode insert depletion issues are virtually eliminated. Though the ECR plasma discharge has been found to typically operate at slightly higher discharge losses than conventional DC ion thrusters (for high total thruster power applications), the discharge power fraction is small (less than 1 percent at 25 kW). In this regard, the benefits of increased life, low discharge plasma potentials, and reduced complexity are welcome tradeoffs for the associated discharge efficiency decrease. Presented here are results from discharge characterization of a large area ECR plasma source for gridded ion thruster applications. These measurements included load matching efficacy, bulk plasma properties via Langmuir probe, and plasma uniformity as measured using current probes distributed at the exit plane. A high degree of plasma uniformity was observed (flatness greater than 0.9). Additionally, charge state composition was qualitatively evaluated using emission spectroscopy. Plasma induced emission was dominated by xenon ion lines. No doubly charged xenon ions were detected.

  6. Charge-exchange plasma generated by an ion thruster

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1977-01-01

    The charge exchange plasma generated by an ion thruster was investigated experimentally using both 5 cm and 15 cm thrusters. Results are shown for wide ranges of radial distance from the thruster and angle from the beam direction. Considerations of test environment, as well as distance from the thruster, indicate that a valid simulation of a thruster on a spacecraft was obtained. A calculation procedure and a sample calculation of charge exchange plasma density and saturation electron current density are included.

  7. Plasma particle simulation of electrostatic ion thrusters

    NASA Technical Reports Server (NTRS)

    Peng, Xiaohang; Keefer, Dennis; Ruyten, Wilhelmus

    1990-01-01

    Charge exchange collisons between beam ions and neutral propellant gas can result in erosion of the accelerator grid surfaces of an ion engine. A particle in cell (PIC) is developed along with a Monte Carlo method to simulate the ion dynamics and charge exchange processes in the grid region of an ion thruster. The simulation is two-dimensional axisymmetric and uses three velocity components (2d3v) to investigate the influence of charge exchange collisions on the ion sputtering of the accelerator grid surfaces. An example calculation has been performed for an ion thruster operated on xenon propellant. The simulation shows that the greatest sputtering occurs on the downstream surface of the grid, but some sputtering can also occur on the upstream surface as well as on the interior of the grid aperture.

  8. Simplified power processing for inert gas ion thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.; Pinero, L. R.; Hamley, J. A.

    1993-01-01

    Significant simplifications to power processors for inert gas ion thrusters in the 1 to 5 kW range have been identified. They include elimination of all but three power supplies - one each for the neutralizer, main discharge, and beam. The neutralizer and discharge power supplies would provide both cathode heating and plasma generating functions. This dual-use power supply concept was validated via integration tests with a 30 cm diameter xenon ion thruster. The beam/accelerator power supply would have positive and negative outputs to allow a single power supply to provide both functions. The discharge and beam power supplies would incorporate full-bridge inverters similar to those proven for flight-ready arcjet propulsion systems. Operation of this simplified power processing scheme at an inverter frequency of 50 kHz results in a projected power processor design with low mass and high efficiency. A 2 kW reference point design has estimated values of specific mass of 5.4 kg/kW and an efficiency of 93 percent.

  9. Models of Plasma Processes in Electrostatic Ion Thrusters

    NASA Astrophysics Data System (ADS)

    Katz, Ira

    2004-11-01

    Under the Project Prometheus Program, NASA is studying the feasibility of large, interplanetary spacecraft using nuclear reactors to provide electricity for multi-kilowatt ion thrusters. For these missions, such as the proposed Jupiter Icy Moons Orbiter (JIMO), ion thrusters will be required to operate for as long as 15 years, several times the longest demonstrated ion thruster life. In order to predict thruster life, a team of researchers at JPL is developing a suite of computer models that describe in two and three dimensions the dominant plasma processes that take place in electrostatic ion xenon thrusters. These models are being developed in close coordination with researchers who measure the plasma properties inside the thruster components. This paper reviews the physics contained in models of plasma generation and current flow in the hollow cathode insert region, discharge chamber ionization processes, grid ion optics, and thruster plumes including primary beam ions, ion-neutral scattering, and charge exchange.

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

    NASA Technical Reports Server (NTRS)

    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.; Anderson, John R.; Bond, Thomas A.; Cardwell, G. I.; Christensen, Jon A.

    1997-01-01

    The NASA Solar Electric Propulsion Technology Applications Readiness Program (NSTAR) will provide a single-string primary propulsion system to NASA's New Millennium Deep Space 1 Mission which will perform comet and asteroid flybys in the years 1999 and 2000. The propulsion system includes a 30-cm diameter ion thruster, a xenon feed system, a power processing unit, and a digital control and interface unit. A total of four engineering model ion thrusters, three breadboard power processors, and a controller have been built, integrated, and tested. An extensive set of development tests has been completed along with thruster design verification tests of 2000 h and 1000 h. An 8000 h Life Demonstration Test is ongoing and has successfully demonstrated more than 6000 h of operation. In situ measurements of accelerator grid wear are consistent with grid lifetimes well in excess of the 12,000 h qualification test requirement. Flight hardware is now being assembled in preparation for integration, functional, and acceptance tests.

  11. High-power and 2.5 kW advanced-technology ion thruster

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.

    1977-01-01

    Investigations for improving ion thruster components in the 30 cm engineering model thruster (EMT) resulted in the demonstration of useful techniques for grid short removal and discharge chamber erosion monitoring, establishment of relationships between double ion production and thruster operating parameters, verification of satisfactory specifications on porous tungsten vaporizer material and barium impregnated porous tungsten inserts, demonstration of a new hollow cathode configuration, and specification of magnetic circuit requirements for reproducing desired magnetic mappings. The capacity of a 30 cm EMT to operate at higher beam voltages and currents (higher power) was determined. Operation at 2 A beam current and higher beam voltage is shown to be essentially equivalent to operation at 1.1 kV with regard to efficiency, lifetime and operating conditions. The only additional requirement is an improvement in high voltage insulation and propellant isolator capacity. Operation at minimum voltage and higher beam currents is shown to increase thruster discharge chamber erosion in proportion to beam current. Studies to find alternatives to molybdenum for manufacturing ion optics grids are also reported.

  12. The 8-CM ion thruster characterization. [mercury ion engine

    NASA Technical Reports Server (NTRS)

    Wessel, F. J.; Williamson, W. S.

    1983-01-01

    The performance capabilities of the 8 cm diameter mercury ion thruster were increased by modifying the thruster operating parameters and component hardware. The initial performance levels, representative of the Hughes/NASA Lewis Research Center Ion Auxiliary Propulsion Subsystem (IAPS) thruster, were raised from the baseline values of thrust, T = 5 mN, and specific impulse, I sub sp = 2,900s, to thrust, T = 25 mN and specific impulse, I sub sp = 4,300 s. Performance characteristics including estmates of the erosion rates of various component surfaces are presented.

  13. NEXT Propellant Management System Integration With Multiple Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Soulas, George C.; Herman, Daniel A.

    2011-01-01

    As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

  14. An 8-cm ion thruster characterization

    NASA Technical Reports Server (NTRS)

    Wessel, F. J.; Hancock, D. J.; Dulgeroff, C. R.; Williamson, W. S.

    1987-01-01

    The performance of the Ion Auxiliary Propulsion System (IAPS) thruster was increased to thrust T = 32 mN, specific impulse I sub sp = 4062 s, and thrust-to-power ratio T/P = 33 mN/kW. This performance was obtained by increasing the discharge power, accelerating voltage, propellant flow rate, and chamber magnetic field. Adding a plenum and main vaporizer for propellant distribution was the only major change required in the thruster. The modified thruster characterization is presented. A cathode magnet assembly did not improve performance. A simplified power processing unit was designed and evaluated. This unit decreased the parts count of the IAPS power processing unit by a factor of ten.

  15. Electron Backstreaming Determination for Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Wirz, Richard E.; Katz, Ira; Goebel, Dan M.; Anderson, John R.

    2008-01-01

    Electron backstreaming in ion thrusters is caused by the random flux of beam electrons past a potential barrier established by the accel grid. A technique that integrates this flux over the radial extent of the barrier reveals important aspects of electron backstreaming phenomena for individual beamlets, across the thruster beam, and throughout thruster life. For individual beamlets it was found that over 99% of the electron backstreaming occurs in a small annulus at the center of the beamlet that is less than 20% the area of the beamlet at the potential barrier established by the accel grid. For the thruster beam it was found that over 99% of the backstreaming current occurs inside of r = 6 cm for the over 28 cm diameter NSTAR grid. Initial validation against ELT data shows that the technique provides the correct behavior and magnitude of electron backstreaming limit, V(sub ebs). From the sensitivity analyses it is apparent that accel grid chamfering may be the dominant mechanism contributing to the sharp rise in the absolute value of V(sub ebs) observed in the ELT but does not explain the rise in ion transparency. Grid gap change also contributes to the absolute value of V(sub ebs) rise and large rises in ion transparency with thruster life for the center gridlet. Screen grid erosion contributes generally to rises in the absolute value of V(sub ebs) and ion transparency, but for the assumptions used herein, it appears to not have as much of an effect chamfering or grid gap change. Overall, it is apparent that accel grid chamfering, grid gap change, and screen grid erosion are important to the increase in electron backstreaming observed during the ELT.

  16. Ion acceleration in electrodeless plasma thrusters

    NASA Astrophysics Data System (ADS)

    Lafleur, Trevor; Cannat, Felix; Jarrige, Julien; Elias, Paul-Quentin; Packan, Denis

    2016-09-01

    Since electrodeless plasma thrusters do not use biased electrodes or grids to accelerate ions, it is unclear what determines the magnitude of the ``accelerating voltage'' and hence what the ion beam energy is. In this work a combined theoretical/experimental study of the relationship between the electron temperature and the ion energy was performed to provide such an answer. Experimental measurements show that the ion energy and electron temperature are strongly correlated, and demonstrate that the driving force for the plasma expansion in magnetic nozzles is the electron pressure: in complete analogy to chemical rockets with physical nozzles. Because there are no electrodes or applied voltages, the plasma that exits the thruster must be current-free, and we show that this establishes a strong criterion that determines the maximum accelerating potential that self-forms in the plasma. This maximum accelerating potential (which is between about 4-6 times the electron temperature) is similar to that which develops for a floating sheath, and depends on the electron velocity distribution function. Based on plasma loss considerations inside the thruster cavity, and the drop-off of the ionization cross section for large electron energies in most gases, we predict a theoretical maximum achievable ion beam energy of about 400 eV for argon and xenon propellants.

  17. Experimental and analytical evaluation of ion thruster/spacecraft interactions

    NASA Technical Reports Server (NTRS)

    Carruth, M. R., Jr. (Editor)

    1981-01-01

    Studies were conducted to both identify the environment produced by ion thrusters and to assess the interaction of this environment on a typical spacecraft and typical science instruments. Spacecraft charging and the charge exchange that accompanies it is discussed in detail. Electromagnetic interference was characterized for ion engines. The electromagnetic compatibility of ion thrusters with spacecraft instruments was determined. The effects of ion thruster plumes on spacecraft were studied with particular emphasis on external surface currents.

  18. Ion Thruster Development at NASA Lewis Research Center

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Hamley, John A.; Patterson, Michael J.; Rawlin, Vincent K.; Sarver-Verhey, Timothy R.

    1992-01-01

    Recent ion propulsion technology efforts at NASA's Lewis Research Center including development of kW-class xenon ion thrusters, high power xenon and krypton ion thrusters, and power processors are reviewed. Thruster physical characteristics, performance data, life projections, and power processor component technology are summarized. The ion propulsion technology program is structured to address a broad set of mission applications from satellite stationkeeping and repositioning to primary propulsion using solar or nuclear power systems.

  19. Numerical Simulation of Ion Thruster Optics

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K. (Technical Monitor); Farnell, Cody C.; Williams, John D.; Wilbur, Paul J.

    2003-01-01

    A three-dimensional simulation code (ffx) designed to analyze ion thruster optics is described. It is an extension of an earlier code and includes special features like the ability to model a wide range of grid geometries, cusp details, and mis-aligned aperture pairs to name a few. However, the principle reason for advancing the code was in the study of ion optics erosion. Ground based testing of ion thruster optics, essential to the understanding of the processes of grid erosion, can be time consuming and costly. Simulation codes that can accurately predict grid lifetimes and the physical mechanisms of grid erosion can be of great utility in the development of future ion thruster optics designed for more ambitious applications. Results of simulations are presented that describe wear profiles for several standard and nonstandard aperture geometries, such as those grid sets with square- or slotted-hole layout patterns. The goal of this paper will be to introduce the methods employed in the ffx code and to briefly demonstrate their use.

  20. 15 cm multipole gas ion thruster

    NASA Technical Reports Server (NTRS)

    Isaacson, G. C.; Kaufman, H. R.

    1976-01-01

    A 15-cm multipole thruster was operated on argon and xenon. The multipole approach used has been shown capable of low discharge losses and flat ion beam profiles with a minimum of redesign. This approach employs low magnetic field strengths and flat or cylindrical sheet-metal parts, hence is suited to rapid optimization and scaling. Only refractory metal cathodes were used in this investigation.

  1. Sputtering erosion in ion and plasma thrusters

    NASA Technical Reports Server (NTRS)

    Ray, Pradosh K.

    1995-01-01

    An experimental set-up to measure low-energy (below 1 keV) sputtering of materials is described. The materials to be bombarded represent ion thruster components as well as insulators used in the stationary plasma thruster. The sputtering takes place in a 9 inch diameter spherical vacuum chamber. Ions of argon, krypton and xenon are used to bombard the target materials. The sputtered neutral atoms are detected by a secondary neutral mass spectrometer (SNMS). Samples of copper, nickel, aluminum, silver and molybdenum are being sputtered initially to calibrate the spectrometer. The base pressure of the chamber is approximately 2 x 10(exp -9) Torr. the primary ion beam is generated by an ion gun which is capable of delivering ion currents in the range of 20 to 500 nA. The ion beam can be focused to a size approximately 1 mm in diameter. The mass spectrometer is positioned 10 mm from the target and at 90 deg angle to the primary ion beam direction. The ion beam impinges on the target at 45 deg. For sputtering of insulators, charge neutralization is performed by flooding the sample with electrons generated from an electron gun. Preliminary sputtering results, methods of calculating the instrument response function of the spectrometer and the relative sensitivity factors of the sputtered elements will be discussed.

  2. Advanced ion thruster and electrochemical launcher research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1983-01-01

    The theoretical model of orificed hollow cathode operation predicted experimentally observed cathode performance with reasonable accuracy. The deflection and divergence characteristics of ion beamlets emanating from a two grid optics system as a function of the relative offset of screen and accel grids hole axes were described. Ion currents associated with discharge chamber operation were controlled to improve ion thruster performance markedly. Limitations imposed by basic physical laws on reductions in screen grid hole size and grid spacing for ion optics systems were described. The influence of stray magnetic fields in the vicinity of a neutralizer on the performance of that neutralizer was demonstrated. The ion current density extracted from a thruster was enhanced by injecting electrons into the region between its ion accelerating grids. Theoretical analysis of the electrothermal ramjet concept of launching space bound payloads at high acceleration levels is described. The operation of this system is broken down into two phases. In the light gas gun phase the payload is accelerated to the velocity at which the ramjet phase can commence. Preliminary models of operation are examined and shown to yield overall energy efficiences for a typical Earth escape launch of 60 to 70%. When shock losses are incorporated these efficiencies are still observed to remain at the relatively high values of 40 to 50%.

  3. NSTAR Ion Thrusters and Power Processors

    NASA Technical Reports Server (NTRS)

    Bond, T. A.; Christensen, J. A.

    1999-01-01

    The purpose of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) project is to validate ion propulsion technology for use on future NASA deep space missions. This program, which was initiated in September 1995, focused on the development of two sets of flight quality ion thrusters, power processors, and controllers that provided the same performance as engineering model hardware and also met the dynamic and environmental requirements of the Deep Space 1 Project. One of the flight sets was used for primary propulsion for the Deep Space 1 spacecraft which was launched in October 1998.

  4. Sputtering Erosion in the Ion Thruster

    NASA Technical Reports Server (NTRS)

    Ray, Pradosh K.; Mantenieks, Maris A. (Technical Monitor)

    2000-01-01

    During the first phase of this research, the sputtering yields of molybdenum by low energy (100 eV and higher) xenon ions were measured by using the methods of secondary neutral mass spectrometry (SNMS) and Rutherford backscattering spectrometry (RBS). However, the measured sputtering yields were found to be far too low to explain the sputtering erosions observed in the long-duration tests of ion thrusters. The only difference between the sputtering yield measurement experiments and the ion thruster tests was that the later are conducted at high ion fluences. Hence, a study was initiated to investigate if any linkage exists between high ion fluence and an enhanced sputtering yield. The objective of this research is to gain an understanding of the causes of the discrepancies between the sputtering rates of molybdenum grids in an ion thruster and those measured from our experiments. We are developing a molecular dynamics simulation technique for studying low-energy xenon ion interactions with molybdenum. It is difficult to determine collision sequences analytically for primary ions below the 200 eV energy range where the ion energy is too low to be able to employ a random cascade model with confidence and it is too high to have to consider only single collision at or near the surface. At these low energies, the range of primary ions is about 1 to 2 nm from the surface and it takes less than 4 collisions on the average to get an ion to degrade to such an energy that it can no longer migrate. The fine details of atomic motion during the sputtering process are revealed through computer simulation schemes. By using an appropriate interatomic potential, the positions and velocities of the incident ion together with a sufficient number of target atoms are determined in small time steps. Hence, it allows one to study the evolution of damages in the target and its effect on the sputtering yield. We are at the preliminary stages of setting up the simulation program.

  5. Erosion rate diagnostics in ion thrusters using laser-induced fluorescence

    NASA Technical Reports Server (NTRS)

    Gaeta, C. J.; Matossian, J. N.; Turley, R. S.; Beattie, J. R.; Williams, J. D.; Williamson, W. S.

    1993-01-01

    We have used laser-induced fluorescence (LIF) to monitor the charge-exchange ion erosion of the molybdenum accelerator electrode in ion thrusters. This real-time, nonintrusive method was implemented by operating a 30cm-diam ring-cusp thruster using xenon propellant. With the thruster operating at a total power of 5 kW, laser radiation at a wavelength of 390 nm (corresponding to a ground state atomic transition of molybdenum) was directed through the extracted ion beam adjacent to the downstream surface of the molybdenum accelerator electrode. Molybdenum atoms, sputtered from this surface as a result of charge-exchange ion erosion, were excited by the laser radiation. The intensity of the laser-induced fluorescence radiation, which is proportional to the sputter rate of the molybdenum atoms, was measured and correlated with variations in thruster operating conditions such as accelerator electrode voltage, accelerator electrode current, and test facility background pressure. We also demonstrated that the LIF technique has sufficient sensitivity and spatial resolution to evaluate accelerator electrode lifetime in ground-based test facilities.

  6. Optical properties of mercury ion thruster exhaust plumes Significance for candidate SEP science instruments. [Solar Electric Propulsion

    NASA Technical Reports Server (NTRS)

    Goldstein, R.; Monahan, K. M.

    1975-01-01

    Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume emission rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter was considered incompatible with simultaneous thruster operation.

  7. The 15 cm mercury ion thruster research 1975

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1975-01-01

    Doubly charged ion current measurements in the beam of a SERT II thruster are shown to introduce corrections which bring its calculated thrust into close agreement with that measured during flight testing. A theoretical model of doubly charged ion production and loss in mercury electron bombardment thrusters is discussed and is shown to yield doubly-to-singly charged ion density ratios that agree with experimental measurements obtained on a 15 cm diameter thruster over a range of operating conditions. Single cusp magnetic field thruster operation is discussed and measured ion beam profiles, performance data, doubly charged ion densities, and discharge plasma characteristics are presented for a range of operating conditions and thruster geometries. Variations in the characteristics of this thruster are compared to those observed in the divergent field thruster and the cusped field thruster is shown to yield flatter ion beam profiles at about the same discharge power and propellant utilization operating point. An ion optics test program is described and the measured effects of grid system dimensions on ion beamlet half angle and diameter are examined. The effectiveness of hollow cathode startup using a thermionically emitting filament within the cathode is examined over a range of mercury flow rates and compared to results obtained with a high voltage tickler startup technique. Results of cathode plasma property measurement tests conducted within the cathode are presented.

  8. Design and Preliminary Testing Plan of Electronegative Ion Thruster

    NASA Technical Reports Server (NTRS)

    Schloeder, Natalie R.; Liu, Thomas M.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.; Aanesland, Ane

    2014-01-01

    Electronegative ion thrusters are a new iteration of existing gridded ion thruster technology differentiated by their ability to produce and accelerate both positive and negative ions. The primary motivations for electronegative ion thruster development include the elimination of lifetime-limiting cathodes from a thruster system and the ability to generate appreciable thrust through the acceleration of both positive or negative-charged ions. Proof-of-concept testing of the PEGASES (Plasma Propulsion with Electronegative GASES) thruster demonstrated the production of positively and negatively-charged ions (argon and sulfur hexafluoride, respectively) in an RF discharge and the subsequent acceleration of each charge species through the application of a time-varying electric field to a pair of metallic grids similar to those found in gridded ion thrusters. Leveraging the knowledge gained through experiments with the PEGASES I and II prototypes, the MINT (Marshall's Ion-ioN Thruster) is being developed to provide a platform for additional electronegative thruster proof-of-concept validation testing including direct thrust measurements. The design criteria used in designing the MINT are outlined and the planned tests that will be used to characterize the performance of the prototype are described.

  9. Ion Thruster Power Levels Extended by a Factor of 10

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.

    2004-01-01

    In response to two NASA Office of Space Science initiatives, the NASA Glenn Research Center is now developing a 7-kW-class xenon ion thruster system for near-term solar-powered spacecraft and a 25-kW ion engine for nuclear-electric spacecraft. The 7-kW ion thruster and power processor can be throttled down to 1 kW and are applicable to 25-kW flagship missions to the outer planets, asteroids, and comets. This propulsion system was scaled up from the 2.5-kW ion thruster and power processor that was developed successfully by Glenn, Boeing, the Jet Propulsion Laboratory (JPL), and Spectrum Astro for the Deep Space 1 spacecraft. The 7-kW ion thruster system is being developed under NASA's Evolutionary Xenon Thruster (NEXT) project, which includes partners from JPL, Aerojet, Boeing, the University of Michigan, and Colorado State University.

  10. Increasing the Life of a Xenon-Ion Spacecraft Thruster

    NASA Technical Reports Server (NTRS)

    Goebel, Dan; Polk, James; Sengupta, Anita; Wirz, Richard

    2007-01-01

    A short document summarizes the redesign of a xenon-ion spacecraft thruster to increase its operational lifetime beyond a limit heretofore imposed by nonuniform ion-impact erosion of an accelerator electrode grid. A peak in the ion current density on the centerline of the thruster causes increased erosion in the center of the grid. The ion-current density in the NSTAR thruster that was the subject of this investigation was characterized by peak-to-average ratio of 2:1 and a peak-to-edge ratio of greater than 10:1. The redesign was directed toward distributing the same beam current more evenly over the entire grid andinvolved several modifications of the magnetic- field topography in the thruster to obtain more nearly uniform ionization. The net result of the redesign was to reduce the peak ion current density by nearly a factor of two, thereby halving the peak erosion rate and doubling the life of the thruster.

  11. Internal erosion rates of a 10-kW xenon ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.

    1988-01-01

    A 30 cm diameter divergent magnetic field ion thruster, developed for mercury operation at 2.7 kW, was modified and operated with xenon propellant at a power level of 10 kW for 567 h to evaluate thruster performance and lifetime. The major differences between this thruster and its baseline configuration were elimination of the three mercury vaporizers, use of a main discharge cathode with a larger orifice, reduction in discharge baffle diameter, and use of an ion accelerating system with larger acceleration grid holes. Grid thickness measurement uncertainties, combined with estimates of the effects of reactive residual facility background gases gave a minimum screen grid lifetime of 7000 h. Discharge cathode orifice erosion rates were measured with three different cathodes with different initial orifice diameters. Three potential problems were identified during the wear test: the upstream side of the discharge baffle eroded at an unacceptable rate; two of the main cathode tubes experienced oxidation, deformation, and failure; and the accelerator grid impingement current was more than an order of magnitude higher than that of the baseline mercury thruster. The charge exchange ion eroison was not quantified in this test. There were no measurable changes in the accelerator grid thickness or the accelerator grid hole diameters.

  12. Internal erosion rates of a 10-kW xenon ion thruster

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.

    1988-01-01

    A 30 cm diameter divergent magnetic field ion thruster, developed for mercury operation at 2.7 kW, was modified and operated with xenon propellant at a power level of 10 kW for 567 h to evaluate thruster performance and lifetime. The major differences between this thruster and its baseline configuration were elimination of the three mercury vaporizers, use of a main discharge cathode with a larger orifice, reduction in discharge baffle diameter, and use of an ion accelerating system with larger acceleration grid holes. Grid thickness measurement uncertainties, combined with estimates of the effects of reactive residual facility background gases gave a minimum screen grid lifetime of 7000 h. Discharge cathode orifice erosion rates were measured with three different cathodes with different initial orifice diameters. Three potential problems were identified during the wear test: the upstream side of the discharge baffle eroded at an unacceptable rate; two of the main cathode tubes experienced oxidation, deformation, and failure; and the accelerator grid impingement current was more than an order of magnitude higher than that of the baseline mercury thruster. The charge exchange ion erosion was not quantified in this test. There were no measurable changes in the accelerator grid thickness or the accelerator grid hole diameters.

  13. Experimental research of radio-frequency ion thruster

    NASA Astrophysics Data System (ADS)

    Antropov, N. N.; Akhmetzhanov, R. V.; Bogatyy, A. V.; Grishin, R. A.; Kozhevnikov, V. V.; Plokhikh, A. P.; Popov, G. A.; Khartov, S. A.

    2016-12-01

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

  14. Initial Thrust Measurements of Marshall's Ion-ioN Thruster

    NASA Technical Reports Server (NTRS)

    Schloeder, Natalie R.; Scogin, Tyler; Liu, Thomas M.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.; Aanesland, Ane

    2015-01-01

    Electronegative ion thrusters are a variation of tradition gridded ion thruster technology differentiated by the production and acceleration of both positive and negative ions. Benefits of electronegative ion thrusters include the elimination of lifetime-limiting cathodes from the thruster architecture and the ability to generate appreciable thrust from both charge species. Following the continued development of electronegative ion thruster technology as exhibited by the PEGASES (Plasma Propulsion with Electronegative GASES) thruster, direct thrust measurements are required to push interest in electronegative ion thruster technology forward. For this work, direct thrust measurements of the MINT (Marshall's Ion-ioN Thruster) will be taken on a hanging pendulum thrust stand for propellant mixtures of Sulfur Hexafluoride and Argon at volumetric flow rates of 5-25 sccm at radio frequency power levels of 100-600 watts at a radio frequency of 13.56 MHz. Acceleration grid operation is operated using a square waveform bias of +/-300 volts at a frequency of 25 kHz.

  15. Next-Generation Ion Thruster Design Tool

    NASA Technical Reports Server (NTRS)

    Stolz, Peter

    2015-01-01

    Computational tools that accurately predict the performance of electric propulsion devices are highly desirable and beneficial to NASA and the broader electric propulsion community. The current state of the art in electric propulsion modeling relies heavily on empirical data and numerous computational "knobs." In Phase I of this project, Tech-X Corporation developed the most detailed ion engine discharge chamber model that currently exists. This kinetic model simulates all particles in the discharge chamber along with a physically correct simulation of the electric fields. In addition, kinetic erosion models are included for modeling the ion-impingement effects on thruster component erosion. In Phase II, Tech-X developed a user-friendly computer program for NASA and other governmental and industry customers. Tech-X has implemented a number of advanced numerical routines to bring the computational time down to a commercially acceptable level. NASA now has a highly sophisticated, user-friendly ion engine discharge chamber modeling tool.

  16. Thermo-mechanical design aspects of mercury bombardment ion thrusters.

    NASA Technical Reports Server (NTRS)

    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.

  17. Evolution of the 1-mlb mercury ion thruster subsystem

    NASA Technical Reports Server (NTRS)

    Kerslake, W. R.; Banks, B. A.

    1978-01-01

    The developmental history, performance, and major lifetests of each component of the present 1-mlb (4.5 mN) thruster system are traced over the past 10 years. The 1-mlb thruster subsystem consists of an 8 cm diameter ion thruster mounted on 2 axis gimbals, a mercury propellant tank, a power electronics unit, a controller/digital interface unit, and necessary electrical harnesses plus propellant tankage and feed lines.

  18. Cusped magnetic field mercury ion thruster. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Beattie, J. R.

    1976-01-01

    The importance of a uniform current density profile in the exhaust beam of an electrostatic ion thruster is discussed in terms of thrust level and accelerator system lifetime. A residence time approach is used to explain the nonuniform beam current density profile of the divergent magnetic field thruster. Mathematical expressions are derived which relate the thruster discharge power loss, propellant utilization, and double to single ion density ratio to the geometry and plasma properties of the discharge chamber. These relationships are applied to a cylindrical discharge chamber model of the thruster. Experimental results are presented for a wide range of the discharge chamber length. The thruster designed for this investigation was operated with a cusped magnetic field as well as a divergent field geometry, and the cusped field geometry is shown to be superior from the standpoint of beam profile uniformity, performance, and double ion population.

  19. Application of the NEXT Ion Thruster Lifetime Assessment to Thruster Throttling

    NASA Technical Reports Server (NTRS)

    VanNoord, Jonathan L.; Herman, Daniel A.

    2010-01-01

    Ion thrusters are low thrust, high specific impulse devices with typical operational lifetimes of 10,000 to 30,000 hr over a range of throttling conditions. The NEXT ion thruster is the latest generation of ion thrusters under development. The NEXT ion thruster currently has a qualification level propellant throughput requirement of 450 kg of xenon, which corresponds to roughly 22,000 hr of operation at the highest input power throttling point. This paper will provide a brief review the previous life assessment predictions for various throttling conditions. A further assessment will be presented examining the anticipated accelerator grid hole wall erosion and related electron backstreaming limit. The continued assessment of the NEXT ion thruster indicates that the first failure mode across the throttling range is expected to be in excess of 36,000 hr of operation from charge exchange induced groove erosion. It is at this duration that the groove is predicted to penetrate the accelerator grid possibly resulting in structural failure. Based on these lifetime and mission assessments, a throttling approach is presented for the Long Duration Test to demonstrate NEXT thruster lifetime and validate modeling.

  20. A high power ion thruster for deep space missions

    NASA Astrophysics Data System (ADS)

    Polk, James E.; Goebel, Dan M.; Snyder, John S.; Schneider, Analyn C.; Johnson, Lee K.; Sengupta, Anita

    2012-07-01

    The Nuclear Electric Xenon Ion System ion thruster was developed for potential outer planet robotic missions using nuclear electric propulsion (NEP). This engine was designed to operate at power levels ranging from 13 to 28 kW at specific impulses of 6000-8500 s and for burn times of up to 10 years. State-of-the-art performance and life assessment tools were used to design the thruster, which featured 57-cm-diameter carbon-carbon composite grids operating at voltages of 3.5-6.5 kV. Preliminary validation of the thruster performance was accomplished with a laboratory model thruster, while in parallel, a flight-like development model (DM) thruster was completed and two DM thrusters fabricated. The first thruster completed full performance testing and a 2000-h wear test. The second successfully completed vibration tests at the full protoflight levels defined for this NEP program and then passed performance validation testing. The thruster design, performance, and the experimental validation of the design tools are discussed in this paper.

  1. High-Power, High-Thrust Ion Thruster (HPHTion)

    NASA Technical Reports Server (NTRS)

    Peterson, Peter Y.

    2015-01-01

    Advances in high-power photovoltaic technology have enabled the possibility of reasonably sized, high-specific power solar arrays. At high specific powers, power levels ranging from 50 to several hundred kilowatts are feasible. Ion thrusters offer long life and overall high efficiency (typically greater than 70 percent efficiency). In Phase I, the team at ElectroDynamic Applications, Inc., built a 25-kW, 50-cm ion thruster discharge chamber and fabricated a laboratory model. This was in response to the need for a single, high-powered engine to fill the gulf between the 7-kW NASA's Evolutionary Xenon Thruster (NEXT) system and a notional 25-kW engine. The Phase II project matured the laboratory model into a protoengineering model ion thruster. This involved the evolution of the discharge chamber to a high-performance thruster by performance testing and characterization via simulated and full beam extraction testing. Through such testing, the team optimized the design and built a protoengineering model thruster. Coupled with gridded ion thruster technology, this technology can enable a wide range of missions, including ambitious near-Earth NASA missions, Department of Defense missions, and commercial satellite activities.

  2. Particle and field measurements on two J-series 30 centimeter thrusters

    NASA Technical Reports Server (NTRS)

    Lathem, W. C.

    1981-01-01

    Tests were performed to characterize the particles and fields associated with two 30 cm mercury ion thrusters operating independently and simultaneously. Flux rates and energies of ions and their distribution around the thrusters were determined. Facility effect ions were measured and their effect on thruster created flux measurements was assessed. The flux rate and distribution of sputtered metal atoms was determined and compared with theory and previous measurements. Mapping of the potential fields in the near vicinity of the thrusters was accomplished.

  3. The High Power Electric Propulsion (HiPEP) Ion Thruster

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Haag, Tom; Patterson, Michael; Williams, George J., Jr.; Sovey, James S.; Carpenter, Christian; Kamhawi, Hani; Malone, Shane; Elliot, Fred

    2004-01-01

    Practical implementation of the proposed Jupiter Icy Moon Orbiter (JIMO) mission, which would require a total delta V of approximately 38 km/s, will require the development of a high power, high specific impulse propulsion system. Initial analyses show that high power gridded ion thrusters could satisfy JIMO mission requirements. A NASA GRC-led team is developing a large area, high specific impulse, nominally 25 kW ion thruster to satisfy both the performance and the lifetime requirements for this proposed mission. The design philosophy and development status as well as a thruster performance assessment are presented.

  4. Ion Beam Characterization of a NEXT Multi-Thruster Array Plume

    NASA Technical Reports Server (NTRS)

    Pencil, Eric J.; Foster, John E.; Patterson, Michael J.; Diaz, Esther M.; Van Noord, Jonathan L.; McEwen, Heather K.

    2006-01-01

    Three operational, engineering model, 7-kW ion thrusters and one instrumented, dormant thruster were installed in a cluster array in a large vacuum facility at NASA Glenn Research Center. A series of engineering demonstration tests were performed to evaluate the system performance impacts of operating various multiple-thruster configurations in an array. A suite of diagnostics was installed to investigate multiple-thruster operation impact on thruster performance and life, thermal interactions, and alternative system modes and architectures. The ion beam characterization included measuring ion current density profiles and ion energy distribution with Faraday probes and retarding potential analyzers, respectively. This report focuses on the ion beam characterization during single thruster operation, multiple thruster operation, various neutralizer configurations, and thruster gimbal articulation. Comparison of beam profiles collected during single and multiple thruster operation demonstrated the utility of superimposing single engine beam profiles to predict multi-thruster beam profiles. High energy ions were detected in the region 45 off the thruster axis, independent of thruster power, number of operating thrusters, and facility background pressure, which indicated that the most probable ion energy was not effected by multiple-thruster operation. There were no significant changes to the beam profiles collected during alternate thruster-neutralizer configurations, therefore supporting the viability of alternative system configuration options. Articulation of one thruster shifted its beam profile, whereas the beam profile of a stationary thruster nearby did not change, indicating there were no beam interactions which was consistent with the behavior of a collisionless beam expansion.

  5. Experimental Results of the Impact of an Ion Thruster Plasma on Microwave Propagation

    NASA Technical Reports Server (NTRS)

    Zaman, Afroz J.; Lambert, Kevin M.

    2000-01-01

    Laboratory at the NASA Glenn Research Center. This facility utilizes a cylindrical, stainless steel, vacuum chamber, which is 18.3 m long and 4.6 m in diameter. For the tests being described here a 30 cm diameter, xenon ion thruster was used. The thruster provided between 500 W and 2.3 kW of operating power. The thruster was mounted on a stand along the axis of the chamber near one of its ends.

  6. Experimental Results of the Impact of an Ion Thruster Plasma on Microwave Propagation

    NASA Technical Reports Server (NTRS)

    Zaman, Afroz J.; Lambert, Kevin M.

    2000-01-01

    Laboratory at the NASA Glenn Research Center. This facility utilizes a cylindrical, stainless steel, vacuum chamber, which is 18.3 m long and 4.6 m in diameter. For the tests being described here a 30 cm diameter, xenon ion thruster was used. The thruster provided between 500 W and 2.3 kW of operating power. The thruster was mounted on a stand along the axis of the chamber near one of its ends and could be moved axially.

  7. An ion thruster internal discharge chamber electrostatic probe diagnostic technique using a high-speed probe positioning system.

    PubMed

    Herman, Daniel A; Gallimore, Alec D

    2008-01-01

    Extensive resources have been allocated to diagnose and minimize lifetime-limiting factors in gridded ion thrusters. While most of this effort has focused on grid erosion, results from wear tests indicate that discharge cathode erosion may also play an important role in limiting the lifetime of ring-cusp ion thrusters proposed for future large flagship missions. The detailed characterization of the near-cathode discharge plasma is essential for mitigating discharge cathode erosion. However, severe difficulty is encountered when attempting to measure internal discharge plasma parameters during thruster operation with conventional probing techniques. These difficulties stem from the high-voltage, high-density discharge cathode plume, which is a hostile environment for probes. A method for interrogating the discharge chamber plasma of a working ion thruster over a two-dimensional grid is demonstrated. The high-speed axial reciprocating probe positioning system is used to minimize thruster perturbation during probe insertion and to reduce heating of the probe. Electrostatic probe measurements from a symmetric double Langmuir probe are presented over a two-dimensional spatial array in the near-discharge cathode assembly region of a 30-cm-diameter ring-cusp ion thruster. Electron temperatures, 2-5 eV, and number density contours, with a maximum of 8 x 10(12) cm(-3) on centerline, are measured. These data provide detailed electron temperature and number density contours which, when combined with plasma potential measurements, may shed light on discharge cathode erosion processes and the effect of thruster operating conditions on erosion rates.

  8. Ion thruster system (8-cm) cyclic endurance test

    NASA Technical Reports Server (NTRS)

    Dulgeroff, C. R.; Beattie, J. R.; Poeschel, R. L.; Hyman, J., Jr.

    1984-01-01

    This report describes the qualification test of an Engineering-Model 5-mN-thrust 8-cm-diameter mercury ion thruster which is representative of the Ion Auxiliary Propulsion System (IAPS) thrusters. Two of these thrusters are scheduled for future flight test. The cyclic endurance test described herein was a ground-based test performed in a vacuum facility with a liquid-nitrogen-cooled cryo-surface and a frozen mercury target. The Power Electronics Unit, Beam Shield, Gimal, and Propellant Tank that were used with the thruster in the endurance test are also similar to those of the IAPS. The IAPS thruster that will undergo the longest beam-on-time during the actual space test will be subjected to 7,055 hours of beam-on-time and 2,557 cycles during the flight test. The endurance test was successfully concluded when the mercury in the IAPS Propellant Tank was consumed. At that time, 8,471 hours of beam-on-time and 599 cycles had been accumulated. Subsequent post-test-evaluation operations were performed (without breaking vacuum) which extended the test values to 652 cycles and 9,489 hours of beam-on-time. The Power Electronic Unit (PEU) and thruster were in the same vacuum chamber throughout the test. The PEU accumulated 10,268 hr of test time with high voltage applied to the operating thruster or dummy load.

  9. Performance Evaluation of the Prototype Model NEXT Ion Thruster

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

    2008-01-01

    The performance testing results of the first prototype model NEXT ion engine, PM1, are presented. The NEXT program has developed the next generation ion propulsion system to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters, with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance. The dry mass of PM1 is 12.7 kg. Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.5 to 6.9 kW. At 6.9 kW, the PM1 thruster demonstrates specific impulse of 4190 s, 237 mN of thrust, and a thrust efficiency of 0.71. The flat beam profile, flatness parameters vary from 0.66 at low-power to 0.88 at full-power, and advanced ion optics reduce localized accelerator grid erosion and increases margins for electron backstreaming, impingement-limited voltage, and screen grid ion transparency. The thruster throughput capability is predicted to exceed 750 kg of xenon, an equivalent of 36,500 hr of continuous operation at the full-power operating condition.

  10. Ion and advanced electric thruster research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1980-01-01

    A phenomenological model of the orificed, hollow cathode based on the field enhanced, thermionic mechanism of electron emission is presented. High frequency oscillations associated with the orificed, hollow cathode are shown to be a consequence of current flow through the cathode orifice. A procedure for Langmuir probing of the hollow cathode discharge and analyzing the resulting probe characteristics is discussed. The results of sputter yield measurements made for molybdenum, tantalum, type 304 stainless steel and copper surfaces being bombarded by low energy argon or mercury ions are also given. The effects of nitrogen and alternated copper layers on the sputter yields of molybdenum, tantalum and 304 stainless steel are also discussed. A dynamic model of electrothermal rocket and ramjet thrusters is developed. The gross performance of these devices is compared to that of an electromagnetic gun for the case of a high acceleration, Earth launch mission. The theoretical performance of electrothermal rockets and ramjets is shown to be comparable to that of the electromagnetic gun.

  11. Double ion production in mercury thrusters. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Peters, R. R.

    1976-01-01

    The development of a model which predicts doubly charged ion density is discussed. The accuracy of the model is shown to be good for two different thruster sizes and a total of 11 different cases. The model indicates that in most cases more than 80% of the doubly charged ions are produced from singly charged ions. This result can be used to develop a much simpler model which, along with correlations of the average plasma properties, can be used to determine the doubly charged ion density in ion thrusters with acceptable accuracy. Two different techniques which can be used to reduce the doubly charged ion density while maintaining good thruster operation, are identified as a result of an examination of the simple model. First, the electron density can be reduced and the thruster size then increased to maintain the same propellant utilization. Second, at a fixed thruster size, the plasma density, temperature and energy can be reduced and then to maintain a constant propellant utilization the open area of the grids to neutral propellant loss can be reduced through the use of a small hole accelerator grid.

  12. Ion Engine and Hall Thruster Development at the NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

    Domonkos, Matthew T.; Patterson, Michael J.; Jankovsky, Robert S.

    2002-01-01

    NASA's Glenn Research Center has been selected to lead development of NASA's Evolutionary Xenon Thruster (NEXT) system. The central feature of the NEXT system is an electric propulsion thruster (EPT) that inherits the knowledge gained through the NSTAR thruster that successfully propelled Deep Space 1 to asteroid Braille and comet Borrelly, while significantly increasing the thruster power level and making improvements in performance parameters associated with NSTAR. The EPT concept under development has a 40 cm beam diameter, twice the effective area of the Deep-Space 1 thruster, while maintaining a relatively-small volume. It incorporates mechanical features and operating conditions to maximize the design heritage established by the flight NSTAR 30 cm engine, while incorporating new technology where warranted to extend the power and throughput capability. The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1.) the development of a laboratory Hall thruster capable of providing high thrust at high power; 2.) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program, These additional activities are related to issues such as thruster lifetime and spacecraft integration.

  13. Initial Thrust Measurements of Marshall's Ion-ioN Thruster

    NASA Technical Reports Server (NTRS)

    Caruso, Natalie R. S.; Scogin, Tyler; Liu, Thomas M.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.

    2015-01-01

    Electronegative ion thrusters are a variation of traditional gridded ion thruster technology differentiated by the production and acceleration of both positive and negative ions. Benefits of electronegative ion thrusters include the elimination of lifetime-limiting cathodes from the thruster architecture and the ability to generate appreciable thrust from both charge species. While much progress has been made in the development of electronegative ion thruster technology, direct thrust measurements are required to unambiguously demonstrate the efficacy of the concept and support continued development. In the present work, direct thrust measurements of the thrust produced by the MINT (Marshall's Ion-ioN Thruster) are performed using an inverted-pendulum thrust stand in the High-Power Electric Propulsion Laboratory's Vacuum Test Facility-1 at the Georgia Institute of Technology with operating pressures ranging from 4.8 x 10(exp -5) and 5.7 x 10(exp -5) torr. Thrust is recorded while operating with a propellant volumetric mixture ratio of 5:1 argon to nitrogen with total volumetric flow rates of 6, 12, and 24 sccm (0.17, 0.34, and 0.68 mg/s). Plasma is generated using a helical antenna at 13.56 MHz and radio frequency (RF) power levels of 150 and 350 W. The acceleration grid assembly is operated using both sinusoidal and square waveform biases of +/-350 V at frequencies of 4, 10, 25, 125, and 225 kHz. Thrust is recorded for two separate thruster configurations: with and without the magnetic filter. No thrust is discernable during thruster operation without the magnetic filter for any volumetric flow rate, RF forward Power level, or acceleration grid biasing scheme. For the full thruster configuration, with the magnetic filter installed, a brief burst of thrust of approximately 3.75 mN +/- 3 mN of error is observed at the start of grid operation for a volumetric flow rate of 24 sccm at 350 W RF power using a sinusoidal waveform grid bias at 125 kHz and +/- 350 V

  14. Prediction of plasma properties in mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Longhurst, G. R.

    1978-01-01

    A simplified theoretical model was developed which obtains to first order the plasma properties in the discharge chamber of a mercury ion thruster from basic thruster design and controllable operating parameters. The basic operation and design of ion thrusters is discussed, and the important processes which influence the plasma properties are described in terms of the design and control parameters. The conservation for mass, charge and energy were applied to the ion production region, which was defined as the region of the discharge chamber having as its outer boundary the surface of revolution of the innermost field line to intersect the anode. Mass conservation and the equations describing the various processes involved with mass addition and removal from the ion production region are satisfied by a Maxwellian electron density spatial distribution in that region.

  15. An Innovative Manufacturing of CCC Ion Thruster Grids by North Carolina A&T's RTM Carbon/Carbon Process

    NASA Technical Reports Server (NTRS)

    Haag, Thomas W. (Technical Monitor); Shivakumar, Kunigal N.

    2003-01-01

    Electric ion thrusters are the preferred engines for deep space missions, because of very high specific impulse. The ion engine consists of screen and accelerator grids containing thousands of concentric very small holes. The xenon gas accelerates between the two grids, thus developing the impulse force. The dominant life-limiting mechanism in the state-of-the-art molybdenum thrusters is the xenon ion sputter erosion of the accelerator grid. Carbon/carbon composites (CCC) have shown to be have less than 1/7 the erosion rates than the molybdenum, thus for interplanetary missions CCC engines are inevitable. Early effort to develop CCC composite thrusters had a limited success because of limitations of the drilling technology and the damage caused by drilling. The proposed is an in-situ manufacturing of holes while the CCC is made. Special low CTE molds will be used along with the NC A&T s patented resin transfer molding (RTM) technology to manufacture the CCC grids. First, a manufacture process for 10-cm diameter thruster grids will be developed and verified. Quality of holes, density, CTE, tension, flexure, transverse fatigue and sputter yield properties will be measured. After establishing the acceptable quality and properties, the process will be scaled to manufacture 30-cm diameter grids. The properties of the two grid sizes are compared with each other.

  16. Ion Thruster Used to Propel the Deep Space 1 Spacecraft to Comet Encounters

    NASA Technical Reports Server (NTRS)

    Sovey, James S.

    2000-01-01

    The NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) Project provided a xenon ion propulsion system to the Deep Space 1 (DS1) spacecraft to validate the propulsion system as well as perform primary propulsion for asteroid and comet encounters. The On-Board Propulsion Branch of the NASA Glenn Research Center at Lewis Field developed engineering model versions of the 30-cm-diameter ion thruster and the 2.5-kW power processor unit (PPU). Glenn then transferred the thruster and PPU technologies to Hughes Electron Dynamics and managed the contract, which supplied two flight sets of thrusters and PPU s to the Deep Space 1 spacecraft and to a ground-based life verification test at the Jet Propulsion Laboratory (JPL). In addition to managing the DS1 spacecraft development, JPL was responsible for the NSTAR Project management, thruster life tests, the feed system, diagnostics, and propulsion subsystem integration. The ion propulsion development team included NASA Glenn, JPL, Hughes Electronics, Moog Inc., and Spectrum Astro Inc. The overall NSTAR subsystem dry mass, including thruster, PPU, controller, cables, and the xenon storage and feed system, is 48 kg. The mass of the xenon stored onboard DS1 was about 81 kg, and the spacecraft wet mass was approximately 500 kg.The DS1 spacecraft was launched on October 24, 1998, and on July 29, 1999, it flew within 16 miles of the small asteroid Braille (formerly 1992KD) at a relative speed of 35,000 mph. As of November 1999, the ion propulsion system had performed flawlessly for nearly 149 days of thrusting. NASA has approved an extension to the mission, which will allow DS1 to continue thrusting to encounters with two comets in 2001. The DS1 optical and plasma diagnostic instruments will be used to investigate the comet and space environments. The spacecraft is scheduled to fly past the dormant comet Wilson- Harrington in January 2001 and the very active comet Borrelly in September 2001, at which time

  17. Thermal Environmental Testing of NSTAR Engineering Model Ion Thrusters

    NASA Technical Reports Server (NTRS)

    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.

  18. Analysis and design of ion thrusters for large space systems

    NASA Technical Reports Server (NTRS)

    James, E. L.

    1980-01-01

    This study undertakes the analysis and conceptual design of a 0.5 Newton electrostatic ion thruster suitable for use on large space system missions in the next decade. Either argon or xenon gas shall be used as propellant. A 50 cm diameter discharge chamber was selected to meet stipulated performance goals. The discharge plasma is contained at the boundary by a periodic structure of alternating permanent magnets generating a series of line cusps. Anode strips between the magnets collect Maxwellian electrons generated by a central cathode. Ion extraction utilizes either two or three grid optics at the user's choice. An extensive analysis was undertaken to investigate optics behavior in the high power environment of this large thruster. A plasma bridge neutralizer operating on inert gas provides charge neutralizing electrons to complete the design. The resulting conceptual thruster and the necessary power management and control requirements are described.

  19. Magnetic field configurations on thruster performance in accordance with ion beam characteristics in cylindrical Hall thruster plasmas

    NASA Astrophysics Data System (ADS)

    Kim, Holak; Choe, Wonho; Lim, Youbong; Lee, Seunghun; Park, Sanghoo

    2017-03-01

    Magnetic field configuration is critical in Hall thrusters for achieving high performance, particularly in thrust, specific impulse, efficiency, etc. Ion beam features are also significantly influenced by magnetic field configurations. In two typical magnetic field configurations (i.e., co-current and counter-current configurations) of a cylindrical Hall thruster, ion beam characteristics are compared in relation to multiply charged ions. Our study shows that the co-current configuration brings about high ion current (or low electron current), high ionization rate, and small plume angle that lead to high thruster performance.

  20. Investigation of Keeper Erosion in the NSTAR Ion Thruster

    NASA Technical Reports Server (NTRS)

    Domonkos, Matthew T.; Foster, John E.; Patterson, Michael J.; Williams, George J., Jr.

    2001-01-01

    The goal of the present investigation was to determine the cause for the difference in the observed discharge keeper erosion between the 8200 hr wear test of a NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) engineering model thruster and the ongoing extended life test (ELT) of the NSTAR flight spare thruster. During the ELT, the NSTAR flight spare ion thruster experienced unanticipated erosion of the discharge cathode keeper. Photographs of the discharge keeper show that the orifice has enlarged to slightly more than twice the original diameter. Several differences between the ELT and the 8200 hr wear test were initially identified to determine any effects which could lead to the erosion in the ELT. In order to identify the cause of the ELT erosion, emission spectra from an engineering model thruster were collected to assess the dependence of keeper erosion on operating conditions. Keeper ion current was measured to estimate wear. Additionally, post-test inspection of both a copper keeper-cap was conducted, and the results are presented. The analysis indicated that the bulk of the ion current was collected within 2-mm radially of the orifice. The estimated volumetric wear in the ELT was comparable to previous wear tests. Redistribution of the ion current on the discharge keeper was determined to be the most likely cause of the ELT erosion. The change in ion current distribution was hypothesized to caused by the modified magnetic field of the flight assemblies.

  1. System for Coupling an IEC Reactor to Ion Thrusters

    NASA Astrophysics Data System (ADS)

    Webber, Jason; Burton, Rodney; Momoto, Hiromu; Miley, George; Richardson, Nathan

    2002-11-01

    A conceptual design for an electric-thruster-driven space ship using a D-He3 fueled Inertial Electrostatic Confinement (IEC) fusion power unit was recently developed [1]. This propulsion system uses a bank of modified NSTAR-type krypton ion thrusters (specific impulse of 16,000 sec.) giving a total thrust of 1020 N. The thrust time for a typical outer planet mission ( e.g. Jupiter) with a delta-V of 50,000 m/s is then 200 days. A key component of this concept is a traveling wave direct energy converter that converts the kinetic energy of 14-MeV fusion reaction product protons to high voltage (about 1 MV) DC electrical output. A unique step-down transformer and rectifier system condition this output for use in the ion thrusters. Details of these components, the NSTAR-thruster modifications plus a magnetic hexa-pole collimator designed to guide the emitted protons into the traveling wave converter will be described. This advanced electric thruster design offers a very high power-to-weight ratio system that is crucial for deep space propulsion. [1] George H. Miley, Hiromu Momota, R. Burton, N.Richardson, M. Coventry, and Y. Shaban, IEC Based D-He3 Fusion for Space Propulsion, Trans Am. Nuclear Society, Annual Meeting, Hollywood, FL, June 2002.

  2. Inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1980-01-01

    Some advances in component technology for inert gas thrusters are described. The maximum electron emission of a hollow cathode with Ar was increased 60-70% by the use of an enclosed keeper configuration. Operation with Ar, but without emissive oxide, was also obtained. A 30 cm thruster operated with Ar at moderate discharge voltages give double-ion measurements consistent with a double ion correlation developed previously using 15 cm thruster data. An attempt was made to reduce discharge losses by biasing anodes positive of the discharge plasma. The reason this attempt was unsuccessful is not yet clear. The performance of a single-grid ion-optics configuration was evaluated. The ion impingement on the single grid accelerator was found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator was 2-3 times the aperture diameter.

  3. NASA's Evolutionary Xenon Thruster (NEXT) Ion Propulsion System Information Summary

    NASA Technical Reports Server (NTRS)

    Pencil, Eirc S.; Benson, Scott W.

    2008-01-01

    This document is a guide to New Frontiers mission proposal teams. The document describes the development and status of the NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system (IPS) technology, its application to planetary missions, and the process anticipated to transition NEXT to the first flight mission.

  4. Cycle life testing of 8-cm mercury ion thruster cathodes

    NASA Technical Reports Server (NTRS)

    Wintucky, E. G.

    1976-01-01

    Two main cathodes have successfully completed 2800 and 1980 cycles and three neutralizers, 3928, 3050, and 2850 cycles in ongoing cycle life tests of flight-type cathode-isolator-vaporizer and neutralizer-isolator-vaporizer assemblies for the 4.45 mN 8-cm Hg ion thruster system. Each cycle included one hour of cathode operation. Starting and operating conditions simulated those expected in a typical auxiliary propulsion mission duty cycle. This paper presents the cycle life test results and also results of an insert comparison test which led to the selection of a rolled foil insert type for the 8-cm Engineering Model Thruster cathodes.

  5. Cycle life testing of 8-cm mercury ion thruster cathodes

    NASA Technical Reports Server (NTRS)

    Wintucky, E. G.

    1976-01-01

    Two main cathodes have successfully completed 2800 and 1980 cycles and three neutralizers, 3928, 3050, and 2850 cycles in ongoing cycle life tests of flight-type cathode-isolator-vaporizer and neutralizer-isolator-vaporizer assemblies for the 4.45 mN 8-cm Hg ion thruster system. Each cycle included one hour of cathode operation. Starting and operating conditions simulated those expected in a typical auxiliary propulsion mission duty cycle. The cycle life test results are presented along with results of an insert comparison test which led to the selection of a rolled foil insert type for the 8-cm Engineering Model Thruster cathodes.

  6. Electric arc discharge damage to ion thruster grids

    NASA Technical Reports Server (NTRS)

    Beebe, D. D.; Nakanishi, S.; Finke, R. C.

    1974-01-01

    Arcs representative of those occurring between the grids of a mercury ion thruster were simulated. Parameters affecting an arc and the resulting damage were studied. The parameters investigated were arc energy, arc duration, and grid geometry. Arc attenuation techniques were also investigated. Potentially serious damage occurred at all energy levels representative of actual thruster operating conditions. Of the grids tested, the lowest open-area configuration sustained the least damage for given conditions. At a fixed energy level a long duration discharge caused greater damage than a short discharge. Attenuation of arc current using various impedances proved to be effective in reducing arc damage. Faults were also deliberately caused using chips of sputtered materials formed during the operation of an actual thruster. These faults were cleared with no serious grid damage resulting using the principles and methods developed in this study.

  7. Ion Propulsion Thruster Including a Plurality of Ion Optic Electrode Pairs

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J. (Inventor)

    2016-01-01

    Ion optics for use in a conventional or annular or other shaped ion thruster are disclosed including a plurality of planar, spaced apart ion optic electrode pairs sized to include a diameter smaller than the diameter of thruster exhaust and retained in, on or otherwise associated with a frame across the thruster exhaust. An electrical connection may be provided for establishing electrical connectivity among a set of first upstream electrodes and an electrical connection may be provided for establishing electrical connectivity among the second downstream electrodes.

  8. A preliminary model of ion beam neutralization. [in thruster plasmas

    NASA Technical Reports Server (NTRS)

    Parks, D. E.; Katz, I.

    1979-01-01

    A theoretical model of neutralized thruster ion beam plasmas has been developed. The basic premise is that the beam forms an electrostatic trap for the neutralizing electrons. A Maxwellian spectrum of electron energies is maintained by collisions between trapped electrons and by collective randomization of velocities of electrons injected from the neutralizer into the surrounding plasma. The theory contains the observed barometric law relationship between electron density and electron temperatures and ion beam spreading in good agreement with measured results.

  9. Magnetic Field Would Reduce Electron Backstreaming in Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Foster, John E.

    2003-01-01

    The imposition of a magnetic field has been proposed as a means of reducing the electron backstreaming problem in ion thrusters. Electron backstreaming refers to the backflow of electrons into the ion thruster. Backstreaming electrons are accelerated by the large potential difference that exists between the ion-thruster acceleration electrodes, which otherwise accelerates positive ions out of the engine to develop thrust. The energetic beam formed by the backstreaming electrons can damage the discharge cathode, as well as other discharge surfaces upstream of the acceleration electrodes. The electron-backstreaming condition occurs when the center potential of the ion accelerator grid is no longer sufficiently negative to prevent electron diffusion back into the ion thruster. This typically occurs over extended periods of operation as accelerator-grid apertures enlarge due to erosion. As a result, ion thrusters are required to operate at increasingly negative accelerator-grid voltages in order to prevent electron backstreaming. These larger negative voltages give rise to higher accelerator grid erosion rates, which in turn accelerates aperture enlargement. Electron backstreaming due to accelerator-gridhole enlargement has been identified as a failure mechanism that will limit ionthruster service lifetime. The proposed method would make it possible to not only reduce the electron backstreaming current at and below the backstreaming voltage limit, but also reduce the backstreaming voltage limit itself. This reduction in the voltage at which electron backstreaming occurs provides operating margin and thereby reduces the magnitude of negative voltage that must be placed on the accelerator grid. Such a reduction reduces accelerator- grid erosion rates. The basic idea behind the proposed method is to impose a spatially uniform magnetic field downstream of the accelerator electrode that is oriented transverse to the thruster axis. The magnetic field must be sufficiently

  10. Cathode-less gridded ion thrusters for small satellites

    NASA Astrophysics Data System (ADS)

    Aanesland, Ane

    2016-10-01

    Electric space propulsion is now a mature technology for commercial satellites and space missions that requires thrust in the order of hundreds of mN, and with available electric power in the order of kW. Developing electric propulsion for SmallSats (1 to 500 kg satellites) are challenging due to the small space and limited available electric power (in the worst case close to 10 W). One of the challenges in downscaling ion and Hall thrusters is the need to neutralize the positive ion beam to prevent beam stalling. This neutralization is achieved by feeding electrons into the downstream space. In most cases hollow cathodes are used for this purpose, but they are fragile and difficult to implement, and in particular for small systems they are difficult to downscale, both in size and electron current. We describe here a new alternative ion thruster that can provide thrust and specific impulse suitable for mission control of satellites as small as 3 kg. The originality of our thruster lies in the acceleration principles and propellant handling. Continuous ion acceleration is achieved by biasing a set of grids with Radio Frequency voltages (RF) via a blocking capacitor. Due to the different mobility of ions and electrons, the blocking capacitor charges up and rectifies the RF voltage. Thus, the ions are accelerated by the self-bias DC voltage. Moreover, due to the RF oscillations, the electrons escape the thruster across the grids during brief instants in the RF period ensuring a full space charge neutralization of the positive ion beam. Due to the RF nature of this system, the space charge limited current increases by almost a factor of 2 compared to classical DC biased grids, which translates into a specific thrust two times higher than for a similar DC system. This new thruster is called Neptune and operates with only one RF power supply for plasma generation, ion acceleration and electron neutralization. We will present the downscaling of this thruster to a 3cm

  11. Modeling Ion Beam Neutralization and Near-Thruster Plume Interactions (POSTPRINT)

    DTIC Science & Technology

    2005-08-31

    charged ions is not desired because it leads to faster thruster erosion. Finally, the thruster was assumed to be a perfect conductor. Electrons absorbed ...July 7-10, 2002 12NASA Glenn Website, NASA’s Evolutionary Xenon Thruster (NEXT), http://space-power.grc.nasa.gov/ ppo /projects/next/accomp.html 13Chen, F

  12. Grid Erosion Modeling of the NEXT Ion Thruster Optics

    NASA Technical Reports Server (NTRS)

    Ernhoff, Jerold W.; Boyd, Iain D.; Soulas, George (Technical Monitor)

    2003-01-01

    Results from several different computational studies of the NEXT ion thruster optics are presented. A study of the effect of beam voltage on accelerator grid aperture wall erosion shows a non-monotonic, complex behavior. Comparison to experimental performance data indicates improvements in simulation of the accelerator grid current, as well as very good agreement with other quantities. Also examined is the effect of ion optics choice on the thruster life, showing that TAG optics provide better margin against electron backstreaming than NSTAR optics. The model is used to predict the change in performance with increasing accelerator grid voltage, showing that although the current collected on the accel grid downstream face increases, the erosion rate decreases. A study is presented for varying doubly-ionized Xenon current fraction. The results show that performance data is not extremely sensitive to the current fraction.

  13. Plans for an in-orbit test of a UK rare gas ion thruster

    SciTech Connect

    Martin, A.R.; Bond, A.; Lavender, K.E.

    1987-01-01

    As part of the developmental work on a UK rare gas ion thruster an exercise has been carried out to provide an initial assessment of the integration of such a thruster into a technology demonstrator satellite. The aim of the mission is to provide an in-orbit flight test of the thruster system, and to evaluate the effect of the thruster ion beam upon the spacecraft and upon the surrounding space plasma. Any interactions with communications links will also be assessed, to check that thruster operation and associated noise levels do not have any unacceptable effect upon the communications. 6 references.

  14. Sputtering Erosion in Ion and Plasma Thrusters

    NASA Technical Reports Server (NTRS)

    Ray, Pradosh K.

    1996-01-01

    Low energy sputtering of molybdenum, tantalum and boron nitride with xenon ions are being studied using secondary neutral and secondary ion mass spectrometry (SNMS/SIMS). An ultrahigh vacuum chamber was used to conduct the experiment at a base pressure of 1x10(exp -9) torr. The primary ion beam is generated by an ion gun which is capable of delivering ion currents in the range of 20 to 500 nA. The ion beam can be focused to a spot size of approximately 1 mm in diameter. The mass spectrometer is positioned 10 mm from the target and 90 deg to the primary ion beam direction. SNMS and SIMS spectra were collected at various incident angles and different ion energies. For boron nitride sputtering, the target was flooded with an electron beam to neutralize the charge buildup on the surface. In the SNMS mode, sputtering of Mo and Ta can be detected at an ion energy as low as 100 eV whereas in boron nitride the same was observed up to an energy of 300 eV. However, in the positive-SIMS mode, the sputtering of Mo was observed at 10 eV incident ion energy. The SIMS spectra obtained for boron nitride clearly identifies the two isotopes of boron as well as cluster ions such as B2(sup +) and molecular ions such as BN(sup +). From the angle versus yields measurements, it was found that the maximum SNMS yield shifts towards lower incident angles at low ion energies for all three samples.

  15. Cathode-less gridded ion thrusters for small satellites

    NASA Astrophysics Data System (ADS)

    Aanesland, Ane; Rafalskyi, Dmytro

    2015-09-01

    We present here a new gridded ion thruster, called Neptune, that operates with only one Radio Frequency (RF) power source for ionization, ion acceleration and beam neutralization in addition to solid iodine as propellant. Thus significant simplifications, over excising gridded thrusters, might allow downscaling to satellites as small as 6 kg. The combined acceleration and neutralization is achieved by applying an RF voltage to the grid system via a blocking capacitor. As for similar RF capacitive systems, a self-bias is formed such that ions are continuously accelerated while electrons are emitted in brief instants within the RF sheath collapse. Moreover, the RF nature of the acceleration system leads to a higher space charge limited current extracted across the grids compared to classical DC operated systems. Measurements of the ion and electron energy distribution functions in the plasma plume show that in addition to the directed beam of ions, the electrons are also anisotropic resulting in a flowing plasma, rather than a beam of positive ions. Experimental characterization of this RF accelerated plume is detailed. This work received financial state aid managed by the ANR as part of the program ``Investissements d'avenir'' under the reference ANR-11-IDEX-0003-02 (Project MINIATURE).

  16. Durability tests of a five centimeter diameter ion thruster system

    NASA Technical Reports Server (NTRS)

    Nakanishi, S.

    1972-01-01

    A modified Hughes SIT-5 system is being tested for durability at the Lewis Research Center. As of October 1, 1972, the thruster subsystem has logged over 8000 hours of operation. The initial 2023 hours were run with a translating screen thrust vector grid. The thruster is currently operating with an electrostatic type vector grid. Profiles and maps taken at widely separated intervals show that performance and operating characteristics have remained essentially constant. Overall efficiency is about 32 percent and power to thrust ratio is 170 watts per millipound at a specific impulse of 2500 seconds. Telescopic examination of the vector grid shows some sputtering erosion due to charge exchange and direct impingement ions.

  17. Microwave ECR Ion Thruster Development Activities at NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Patterson, Michael J.

    2002-01-01

    Outer solar system missions will have propulsion system lifetime requirements well in excess of that which can be satisfied by ion thrusters utilizing conventional hollow cathode technology. To satisfy such mission requirements, other technologies must be investigated. One possible approach is to utilize electrodeless plasma production schemes. Such an approach has seen low power application less than 1 kW on earth-space spacecraft such as ARTEMIS which uses the rf thruster the RIT 10 and deep space missions such as MUSES-C which will use a microwave ion thruster. Microwave and rf thruster technologies are compared. A microwave-based ion thruster is investigated for potential high power ion thruster systems requiring very long lifetimes.

  18. TADPOLE satellite. [low cost synchronous orbit satellite to evaluate small mercury bombardment ion thruster applications

    NASA Technical Reports Server (NTRS)

    1974-01-01

    A low cost synchronous orbit satellite to evaluate small mercury bombardment ion thruster applications is described. The ion thrusters provide the satellite with precise north-south and east-west stationkeeping capabilities. In addition, the thrusters are used to unload the reaction wheels used for attitude control and for other purposes described in the report. The proposed satellite is named TADPOLE. (Technology Application Demonstration Program of Low Energy).

  19. Inert-gas thruster technology

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.; Trock, D. C.

    1981-01-01

    Attention is given to recent advances in component technology for inert-gas thrusters. It is noted that the maximum electron emission of a hollow cathode with Ar can be increased 60-70% by using an enclosed keeper configuration. Operation with Ar but without emissive oxide has also been attained. A 30-cm thruster operated with Ar at moderate discharge voltages is found to give double-ion measurements consistent with a double-ion correlation developed earlier on the basis of 15-cm thruster data. An attempt is made to reduce discharge losses by biasing anodes positive of the discharge plasma. The performance of a single-grid ion-optics configuration is assessed. The ion impingement on the single-grid accelerator is found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator is 2-3 times the aperture diameter.

  20. Tailoring Ion-Thruster Grid Apertures for Greater Efficiency

    NASA Technical Reports Server (NTRS)

    Brophy, John

    2004-01-01

    A report proposes tailoring the diameters of the apertures in the accelerator grid of an ion thruster to reduce the open grid area through which un-ionized propellant gas can escape. The result would be a reduction in the loss of propellant gas and a corresponding increase in propellant efficiency. In a typical ion thruster, the plasma density decreases with radial distance from the centerline, and as a consequence, the diameters of ion beamlets decrease with increasing radial distance. According to the proposal, the apertures, through which the ion beamlets must pass, would be sized to match the diameters (with margin) of the beamlets. The decrease of the aperture diameters with radial distance would result in a significant reduction in the open grid area: In an example based on representative design parameters, the reduction could be as much as 30 percent. In this example, the transparency to un-ionized propellant would decrease from 0.24 to 0.17 and, as a result, the propellant efficiency would increase from 0.91 to 0.96.

  1. Qualification of Commercial XIPS(R) Ion Thrusters for NASA Deep Space Missions

    NASA Technical Reports Server (NTRS)

    Goebel, Dan M.; Polk, James E.; Wirz, Richard E.; Snyder, J.Steven; Mikellides, Ioannis G.; Katz, Ira; Anderson, John

    2008-01-01

    Electric propulsion systems based on commercial ion and Hall thrusters have the potential for significantly reducing the cost and schedule-risk of Ion Propulsion Systems (IPS) for deep space missions. The large fleet of geosynchronous communication satellites that use solar electric propulsion (SEP), which will approach 40 satellites by year-end, demonstrates the significant level of technical maturity and spaceflight heritage achieved by the commercial IPS systems. A program to delta-qualify XIPS(R) ion thrusters for deep space missions is underway at JPL. This program includes modeling of the thruster grid and cathode life, environmental testing of a 25-centimeter electromagnetic (EM) thruster over DAWN-like vibe and temperature profiles, and wear testing of the thruster cathodes to demonstrate the life and benchmark the model results. This paper will present the delta-qualification status of the XIPS thruster and discuss the life and reliability with respect to known failure mechanisms.

  2. Magnetic shielding of walls from the unmagnetized ion beam in a Hall thruster

    SciTech Connect

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

    2013-01-14

    We demonstrate by numerical simulations and experiments that the unmagnetized ion beam formed in a Hall thruster can be controlled by an applied magnetic field in a manner that reduces by 2-3 orders of magnitude deleterious ion bombardment of the containing walls. The suppression of wall erosion in Hall thrusters to such low levels has remained elusive for decades.

  3. Ring-cusp ion thruster with shell anode

    NASA Technical Reports Server (NTRS)

    Sovey, J. S.; Rawlin, V. K.; Roman, R. F. (Inventor)

    1984-01-01

    An improved ion thruster for low specific impulse operation in the 1500 sec to 6000 sec range has a multicusp boundary field provided by high strength magnets on an iron anode shell which lengthens the paths of electrons from a hollow cathode assembly. A downstream anode pole piece in the form of an iron ring supports a ring of magnets to provide a more uniform beam profile. A cylindrical cathode magnet can be moved selectively in an axial direction along a feed tube to produce the desired magnetic field at the cathode tip.

  4. Simplification of power electronics for ion thruster neutralizers

    NASA Technical Reports Server (NTRS)

    Gruber, R. P.

    1982-01-01

    A need exists for less complex and lower cost ion thruster systems. Design approaches and the demonstration of neutralizer power electronics for relaxed neutralizer keeper, tip heater, and vaporizer requirements are discussed. The neutralizer circuitry is operated from a 200 to 400 V bus and demonstrates an order of magnitude reduction in parts count. Furthermore, a new technique is described for regulating tip heater power and automatically switching over to provide keeper power with only four additional components. A new design to control the flow rate of the neutralizer with one integrated circuit is also presented.

  5. Simplified Ion Thruster Xenon Feed System for NASA Science Missions

    NASA Technical Reports Server (NTRS)

    Snyder, John Steven; Randolph, Thomas M.; Hofer, Richard R.; Goebel, Dan M.

    2009-01-01

    The successful implementation of ion thruster technology on the Deep Space 1 technology demonstration mission paved the way for its first use on the Dawn science mission, which launched in September 2007. Both Deep Space 1 and Dawn used a "bang-bang" xenon feed system which has proven to be highly successful. This type of feed system, however, is complex with many parts and requires a significant amount of engineering work for architecture changes. A simplified feed system, with fewer parts and less engineering work for architecture changes, is desirable to reduce the feed system cost to future missions. An attractive new path for ion thruster feed systems is based on new components developed by industry in support of commercial applications of electric propulsion systems. For example, since the launch of Deep Space 1 tens of mechanical xenon pressure regulators have successfully flown on commercial spacecraft using electric propulsion. In addition, active proportional flow controllers have flown on the Hall-thruster-equipped Tacsat-2, are flying on the ion thruster GOCE mission, and will fly next year on the Advanced EHF spacecraft. This present paper briefly reviews the Dawn xenon feed system and those implemented on other xenon electric propulsion flight missions. A simplified feed system architecture is presented that is based on assembling flight-qualified components in a manner that will reduce non-recurring engineering associated with propulsion system architecture changes, and is compared to the NASA Dawn standard. The simplified feed system includes, compared to Dawn, passive high-pressure regulation, a reduced part count, reduced complexity due to cross-strapping, and reduced non-recurring engineering work required for feed system changes. A demonstration feed system was assembled using flight-like components and used to operate a laboratory NSTAR-class ion engine. Feed system components integrated into a single-string architecture successfully operated

  6. Ion angular distribution simulation of the HEMP Thruster

    NASA Astrophysics Data System (ADS)

    Duras, Julia; Koch, Norbert; Kahnfeld, Daniel; Bandelow, Gunnar; Matthias, Paul; Lüskow, Karl Felix; Schneider, Ralf; Kemnitz, Stefan

    2016-10-01

    Ion angular current and energy distributions are important parameters for ion thrusters, which are typically measured at a few tens of centimetres to a few meters distance from thruster exit. However, fully kinetic Particle-in-Cell simulations are not able to simulate such domain sizes, due to high computational costs. Therefore, a parallelisation strategy of the code is presented to reduce computational time. To map diagnostics information from the domain boundary of the calculational domain to the positions of experimental diagnostics the concept of transfer functions is introduced. The calculated ion beam angular distributions in the plume region are quite sensitive to boundary conditions of the potential, possible additional source contributions, e.g. from secondary electron emission at vessel walls, and charge exchange collisions. This work was supported by the Bavarian State Ministry of Education Science and the Arts and the German Space Agency DLR. We also like to thank R. Heidemann from THALES Electron Devices GmbH, for interesting and stimulating discussions.

  7. The PEGASES gridded ion-ion thruster physics, performance and predictions

    NASA Astrophysics Data System (ADS)

    Aanesland, Ane; Rafalskyi, Dmytro; Bredin, Jerome; Grondein, Pascaline; Oudini, Noureddine; Chabert, Pascal

    2013-09-01

    The PEGASES (Plasma propulsion with Electronegative gases) thruster is a gridded ion thruster that accelerates alternately positively and negatively charged ions to provide thrust. Over the last years various prototypes have been tested, adequate diagnostics have been developed and analytical models and simulations are made to better understand and control the physics involved. The plasma density in the region of the ion-ion plasma predicts that the performance of the PEGASES thruster can be comparable to existing thrusters on the market. We have recently provided the first experimental proof-of-concept, accelerating alternately positive and negative ions from an ion-ion plasma within a 10 kHz cycle. Here we present the state of the art in the PEGASES development and discuss the various physics involved and its possible future in space. This work is funded by EADS Astrium, ANR (Agence nationale de la recherche) under contract ANR-11-BS09-040 and FP7 under contract PIIF-GA-2012-326054.

  8. Theoretical investigations of plasma processes in the ion bombardment thruster

    NASA Technical Reports Server (NTRS)

    Wilhelm, H. E.

    1975-01-01

    A physical model for a thruster discharge was developed, consisting of a spatially diverging plasma sustained electrically between a small ring cathode and a larger ring anode in a cylindrical chamber with an axial magnetic field. The associated boundary-value problem for the coupled partial differential equations with mixed boundary conditions, which describe the electric potential and the plasma velocity fields, was solved in closed form. By means of quantum-mechanical perturbation theory, a formula for the number S(E) of atoms sputtered on the average by an ion of energy E was derived from first principles. The boundary-value problem describing the diffusion of the sputtered atoms through the surrounding rarefied electron-ion plasma to the system surfaces of ion propulsion systems was formulated and treated analytically. It is shown that outer boundary-value problems of this type lead to a complex integral equation, which requires numerical resolution.

  9. The effects of an ion-thruster exhaust plume on S-band carrier transmission

    NASA Technical Reports Server (NTRS)

    Ackerknecht, W. E.; Stanton, P. H.

    1976-01-01

    The study reported here was undertaken (1) to develop models of the effects of an ion-thruster exhaust plume on S-band signals, and (2) to measure the effects. The results show that an S-band signal passing through an ion-thruster plume is reduced in amplitude and advanced in phase. The mathematical models gave reasonable estimates of the average signal attenuation and phase shift. Negligible fluctuations in the signal amplitude and phase were measured during steady-state thruster operation. However, large jumps in phase occurred when changes were made in the thruster operating state. This study confirms that the thruster plume can have a significant effect on S-band communication link performance; hence the plume effects must be considered in S-band link calculations when electric thrusters are used for spacecraft propulsion.

  10. Xenon Sputter Yield Measurements for Ion Thruster Materials

    NASA Technical Reports Server (NTRS)

    Williams, John D.; Gardner, Michael M.; Johnson, Mark L.; Wilbur, Paul J.

    2003-01-01

    In this paper, we describe a technique that was used to measure total and differential sputter yields of materials important to high specific impulse ion thrusters. The heart of the technique is a quartz crystal monitor that is swept at constant radial distance from a small target region where a high current density xenon ion beam is aimed. Differential sputtering yields were generally measured over a full 180 deg arc in a plane that included the beam centerline and the normal vector to the target surface. Sputter yield results are presented for a xenon ion energy range from 0.5 to 10 keV and an angle of incidence range from 0 deg to 70 deg from the target surface normal direction for targets consisting of molybdenum, titanium, solid (Poco) graphite, and flexible graphite (grafoil). Total sputter yields are calculated using a simple integration procedure and comparisons are made to sputter yields obtained from the literature. In general, the agreement between the available data is good. As expected for heavy xenon ions, the differential and total sputter yields are found to be strong functions of angle of incidence. Significant under- and over-cosine behavior is observed at low- and high-ion energies, respectively. In addition, strong differences in differential yield behavior are observed between low-Z targets (C and Ti) and high-Z targets (Mo). Curve fits to the differential sputter yield data are provided. They should prove useful to analysts interested in predicting the erosion profiles of ion thruster components and determining where the erosion products re-deposit.

  11. Development Efforts Expanded in Ion Propulsion: Ion Thrusters Developed With Higher Power Levels

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Rawlin, Vincent K.; Sovey, James S.

    2003-01-01

    The NASA Glenn Research Center was the major contributor of 2-kW-class ion thruster technology to the Deep Space 1 mission, which was successfully completed in early 2002. Recently, NASA s Office of Space Science awarded approximately $21 million to Glenn to develop higher power xenon ion propulsion systems for large flagship missions such as outer planet explorers and sample return missions. The project, referred to as NASA's Evolutionary Xenon Thruster (NEXT), is a logical follow-on to the ion propulsion system demonstrated on Deep Space 1. The propulsion system power level for NEXT is expected to be as high as 25 kW, incorporating multiple ion thrusters, each capable of being throttled over a 1- to 6-kW power range. To date, engineering model thrusters have been developed, and performance and plume diagnostics are now being documented. The project team-Glenn, the Jet Propulsion Laboratory, General Dynamics, Boeing Electron Dynamic Devices, the Applied Physics Laboratory, the University of Michigan, and Colorado State University-is in the process of developing hardware for a ground demonstration of the NEXT propulsion system, which comprises a xenon feed system, controllers, multiple thrusters, and power processors. The development program also will include life assessments by tests and analyses, single-string tests of ion thrusters and power systems, and finally, multistring thruster system tests in calendar year 2005. In addition, NASA's Office of Space Science selected Glenn to lead the development of a 25-kW xenon thruster to enable NASA to conduct future missions to the outer planets of Jupiter and beyond, under the High Power Electric Propulsion (HiPEP) program. The development of a 100-kW-class ion propulsion system and power conversion systems are critical components to enable future nuclear-electric propulsion systems. In fiscal year 2003, a team composed of Glenn, the Boeing Company, General Dynamics, the Applied Physics Laboratory, the Naval Research

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

    NASA Technical Reports Server (NTRS)

    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.

  13. Design of an ion thruster movable grid thrust vectoring system

    NASA Astrophysics Data System (ADS)

    Kural, Aleksander; Leveque, Nicolas; Welch, Chris; Wolanski, Piotr

    2004-08-01

    Several reasons justify the development of an ion propulsion system thrust vectoring system. Spacecraft launched to date have used ion thrusters mounted on gimbals to control the thrust vector within a range of about ±5°. Such devices have large mass and dimensions, hence the need exists for a more compact system, preferably mounted within the thruster itself. Since the 1970s several thrust vectoring systems have been developed, with the translatable accelerator grid electrode being considered the most promising. Laboratory models of this system have already been built and successfully tested, but there is still room for improvement in their mechanical design. This work aims to investigate possibilities of refining the design of such movable grid thrust vectoring systems. Two grid suspension designs and three types of actuators were evaluated. The actuators examined were a micro electromechanical system, a NanoMuscle shape memory alloy actuator and a piezoelectric driver. Criteria used for choosing the best system included mechanical simplicity (use of the fewest mechanical parts), accuracy, power consumption and behaviour in space conditions. Designs of systems using these actuators are proposed. In addition, a mission to Mercury using the system with piezoelectric drivers has been modelled and its performance presented.

  14. High performance auxiliary-propulsion ion thruster with ion-machined accelerator grid

    NASA Technical Reports Server (NTRS)

    Hudson, W. R.; Banks, B. A.

    1975-01-01

    An improvement in thruster performance was achieved by reducing the diameter of the accelerator grid holes. The smaller accelerator grid holes resulted in a reduction in neutral mercury atoms escaping the discharge chamber, which in turn enhanced the discharge propellant utilization from approximately 68 percent to 92 percent. The accelerator grids were fabricated by ion machining with an 8-centimeter-diameter thruster, and the screen grid holes individually focused ion beamlets onto the blank accelerator grid. The resulting accelerator grid holes are less than 1.12 millimeters in diameter, while previously used accelerator grids had hole diameters of 1.69 millimeters. The thruster could be operated with the small-hole accelerator grid at neutralizer potential.

  15. Discharge Hollow Cathode and Extraction Grid Analysis for the MiXI Ion Thruster

    NASA Technical Reports Server (NTRS)

    Wirz, Richard; Sullivan, Regina; Przybylowski, JoHanna; Silva, Mike

    2006-01-01

    Miniature ion thrusters are well-suited future space missions such as Terrestrial Planet Finder - Interferometer (TPF-I), where high efficiency thrusters using non-contaminating noble gas propellant are desirable. Transient dynamic and orbital analyses have shown that the low-noise, continuous thrust of the Miniature Xenon Ion (MiXI) thruster is desirable for TPF-I formation rotation maneuvers when compared with other thruster options [1], [2]. The 3cm diameter MiXI thruster, Figure 1, was originally designed using experimental methods and is capable of high Isp (> 3,000 sec), propellant efficiency > 80%, and thrust from <0.1 mN to >1.5 mN [3]. The MiXI thruster must demonstrate high levels of thrust resolution and a low minimum impulse bit to ensure it meets the precision formation flying needs of missions such as TPF-I. A novel concept for controlling the ion extraction voltages yields the necessary thrust characteristics for the MiXI thruster. Experiments verify these techniques and two dimensional computational models show that such techniques should have minimal effect on the lifetime of the thruster. During this effort, the MiXI thruster incorporates, for the first time, flight like hollow cathodes for both the discharge chamber and beam neutralization.

  16. Performance optimized, small structurally integrated ion thruster system

    NASA Technical Reports Server (NTRS)

    Hyman, J., Jr.

    1973-01-01

    A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.

  17. Nonlinear ion dynamics in Hall thruster plasma source by ion transit-time instability

    NASA Astrophysics Data System (ADS)

    Lim, Youbong; Choe, Wonho; Mazouffre, Stéphane; Park, Jae Sun; Kim, Holak; Seon, Jongho; Garrigues, L.

    2017-03-01

    High-energy tail formation in an ion energy distribution function (IEDF) is explained in a Hall thruster plasma with the stationary crossed electric and magnetic fields whose discharge current is oscillated at the ion transit-time scale with a frequency of 360 kHz. Among ions in different charge states, singly charged Xe ions (Xe+) have an IEDF that is significantly broadened and shifted toward the high-energy side, which contributes to tail formation in the entire IEDF. Analytical and numerical investigations confirm that the IEDF tail is due to nonlinear ion dynamics in the ion transit-time oscillation.

  18. Wear characteristics from the extended lift test of the DS1 flight spare ion thruster

    NASA Technical Reports Server (NTRS)

    Sengupta, A.; Brophy, J. R.; Goodfellow, K. D.

    2003-01-01

    An on-going life test of the Deep Space One flight spare ion thruster, is being conducted at the Jet Propulsion Laboratory. The thruster has operated for over 27,290 hours and processed in excess of 219 kg of xenon propellant.

  19. Modeling Ion Beam Neutralization and Near-Thruster Plume Interactions

    DTIC Science & Technology

    2005-08-31

    desired because it leads to faster thruster erosion. Finally, the thruster was assumed to be a perfect conductor. Any absorbed electrons were re...NASA’s Evolutionary Xenon Thruster (NEXT), http://space-power.grc.nasa.gov/ ppo /projects/next/accomp.html 13Chen, F. F., Introduction to Plasma Physics

  20. Plasma generation near an Ion thruster disharge chamber hollow cathode

    NASA Technical Reports Server (NTRS)

    Katz, Ira; Anderson, John R.; Goebel, Dan M.; Wirz, Richard; Sengupta, Anita

    2003-01-01

    In gridded electrostatic thrusters, ions are produced by electron bombardment in the discharge chamber. In most of these thrusters, a single, centrally located hollow cathode supplies the ionizing electrons. An applied magnetic field in the discharge chamber restricts the electrons leaving the hollow cathode to a very narrow channel. In this channel, the high electron current density ionizes both propellant gas flowing from the hollow cathode, and other neutrals from the main propellant flow from the plenum. The processes that occur just past the hollow cathode exit are very important. In recent engine tests, several cases of discharge cathode orifice place and keeper erosion have been reported. In this paper we present results from a new 1-D, variable area model of the plasma processes in the magnetized channel just downstream of the hollow cathode keeper. The model predicts plasma densities, and temperatures consistent with those reported in the literature for the NSTAR engine, and preliminary results from the model show a potential maximum just downstream of the cathode.

  1. A north-south stationkeeping ion thruster system for ATS-F.

    NASA Technical Reports Server (NTRS)

    Worlock, R.; James, E.; Ramsey, W.; Trump, G.; Gant, G.; Jan, L.; Bartlett, R.

    1972-01-01

    An ion thruster system is being developed for the ATS-F satellite to demonstrate the application of ion thruster technology to the synchronous satellite north-south stationkeeping mission. The cesium bombardment ion thruster develops one millipound thrust at 2600 seconds specific impulse and provides thrust vectoring by accelerator electrode displacement. The propellant system is sized for two years operation at 25 percent duty cycle. Power conditioning circuitry is based on transistor inverters switching at 10 kHz. Thirteen command channels allow flexibility in operation; 12 telemetry channels provide information on system performance. Input power is less than 150 watts.

  2. Cylindrical Langmuir probe measurements in an ion thruster plume

    SciTech Connect

    Semak, V.V.; Keefer, D.

    1995-12-31

    Cylindrical Langmuir probe measurements have been made in the exhaust plume of a 15 cm diameter three-grid ion thruster. Xenon gas was used as a propellant. The experiments were performed in a large vacuum chamber with background pressure values of 2 {times} 10{sup {minus}4} Torr and 3 {times} 10{sup {minus}6} Torr. A 250 {micro}m diameter tungsten probe 2.78 mm long was used in the experiments. The data collected for radial and axial scans were analyzed to provide measurements of electron temperature, plasma potential and electron density. It was found that, unlike the case of a stationary plasma, the electron saturation current is only several times higher than ion current. Small errors in the determination of the ion current component of the characteristic curve can result in significant overestimation of electron temperature. A method of data analysis was developed which includes an estimation of the collected ion current. This method is compared with a numerical particle simulation model for calculation of the ion current component.

  3. Electric field measurement in microwave discharge ion thruster with electro-optic probe.

    PubMed

    Ise, Toshiyuki; Tsukizaki, Ryudo; Togo, Hiroyoshi; Koizumi, Hiroyuki; Kuninaka, Hitoshi

    2012-12-01

    In order to understand the internal phenomena in a microwave discharge ion thruster, it is important to measure the distribution of the microwave electric field inside the discharge chamber, which is directly related to the plasma production. In this study, we proposed a novel method of measuring a microwave electric field with an electro-optic (EO) probe based on the Pockels effect. The probe, including a cooling system, contains no metal and can be accessed in the discharge chamber with less disruption to the microwave distribution. This method enables measurement of the electric field profile under ion beam acceleration. We first verified the measurement with the EO probe by a comparison with a finite-difference time domain numerical simulation of the microwave electric field in atmosphere. Second, we showed that the deviations of the reflected microwave power and the beam current were less than 8% due to inserting the EO probe into the ion thruster under ion beam acceleration. Finally, we successfully demonstrated the measurement of the electric-field profile in the ion thruster under ion beam acceleration. These measurements show that the electric field distribution in the thruster dramatically changes in the ion thruster under ion beam acceleration as the propellant mass flow rate increases. These results indicate that this new method using an EO probe can provide a useful guide for improving the propulsion of microwave discharge ion thrusters.

  4. Electric field measurement in microwave discharge ion thruster with electro-optic probe

    NASA Astrophysics Data System (ADS)

    Ise, Toshiyuki; Tsukizaki, Ryudo; Togo, Hiroyoshi; Koizumi, Hiroyuki; Kuninaka, Hitoshi

    2012-12-01

    In order to understand the internal phenomena in a microwave discharge ion thruster, it is important to measure the distribution of the microwave electric field inside the discharge chamber, which is directly related to the plasma production. In this study, we proposed a novel method of measuring a microwave electric field with an electro-optic (EO) probe based on the Pockels effect. The probe, including a cooling system, contains no metal and can be accessed in the discharge chamber with less disruption to the microwave distribution. This method enables measurement of the electric field profile under ion beam acceleration. We first verified the measurement with the EO probe by a comparison with a finite-difference time domain numerical simulation of the microwave electric field in atmosphere. Second, we showed that the deviations of the reflected microwave power and the beam current were less than 8% due to inserting the EO probe into the ion thruster under ion beam acceleration. Finally, we successfully demonstrated the measurement of the electric-field profile in the ion thruster under ion beam acceleration. These measurements show that the electric field distribution in the thruster dramatically changes in the ion thruster under ion beam acceleration as the propellant mass flow rate increases. These results indicate that this new method using an EO probe can provide a useful guide for improving the propulsion of microwave discharge ion thrusters.

  5. Performance Evaluation of an Expanded Range XIPS Ion Thruster System for NASA Science Missions

    NASA Technical Reports Server (NTRS)

    Oh, David Y.; Goebel, Dan M.

    2006-01-01

    This paper examines the benefit that a solar electric propulsion (SEP) system based on the 5 kW Xenon Ion Propulsion System (XIPS) could have for NASA's Discovery class deep space missions. The relative cost and performance of the commercial heritage XIPS system is compared to NSTAR ion thruster based systems on three Discovery class reference missions: 1) a Near Earth Asteroid Sample Return, 2) a Comet Rendezvous and 3) a Main Belt Asteroid Rendezvous. It is found that systems utilizing a single operating XIPS thruster provides significant performance advantages over a single operating NSTAR thruster. In fact, XIPS performs as well as systems utilizing two operating NSTAR thrusters, and still costs less than the NSTAR system with a single operating thruster. This makes XIPS based SEP a competitive and attractive candidate for Discovery class science missions.

  6. In Situ, Time-Resolved Accelerator Grid Erosion Measurements in the NSTAR 8000 Hour Ion Engine Wear Test

    NASA Technical Reports Server (NTRS)

    Sovey, J.

    1997-01-01

    Time-resolved, in situ measurements of the charge exchange ion erosion pattern on the downstream face of the accelerator grid have been made during an ongoin wear test of the NSTAR 30 cm ion thruster.

  7. Charge-exchange plasma generated by an ion thruster

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1975-01-01

    The use of high voltage solar arrays greatly reduces or eliminates power processing requirements in space electric propulsion systems. This use also requires substantial areas of solar array to be at high positive potential relative to space and most of the spacecraft. The charge exchange plasma conducts electrons from the ion beam to such positive surfaces, and thereby electrically load the high voltage solar array. To evaluate this problem, the charge-exchange plasma generated by an ion beam was investigated experimentally. Based upon the experimental data, a simple model was derived for the charge-exchange plasma. This model is conservative in the sense that both the electron/ion density and the electron current density should be equal to, or less than, the preducted value for all directions in the hemisphere upstream of the ion beam direction. Increasing the distance between a positive potential surface (such as a high voltage solar array) and the thruster is the simplest way to control interactions. Both densities and currents vary as the inverse square of this distance.

  8. Domed, 40-cm-Diameter Ion Optics for an Ion Thruster

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Haag, Thomas W.; Patterson, Michael J.

    2006-01-01

    Improved accelerator and screen grids for an ion accelerator have been designed and tested in a continuing effort to increase the sustainable power and thrust at the high end of the accelerator throttling range. The accelerator and screen grids are undergoing development for intended use as NASA s Evolutionary Xenon Thruster (NEXT) a spacecraft thruster that would have an input-power throttling range of 1.2 to 6.9 kW. The improved accelerator and screen grids could also be incorporated into ion accelerators used in such industrial processes as ion implantation and ion milling. NEXT is a successor to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster - a state-of-the-art ion thruster characterized by, among other things, a beam-extraction diameter of 28 cm, a span-to-gap ratio (defined as this diameter divided by the distance between the grids) of about 430, and a rated peak input power of 2.3 kW. To enable the NEXT thruster to operate at the required higher peak power, the beam-extraction diameter was increased to 40 cm almost doubling the beam-extraction area over that of NSTAR (see figure). The span-to-gap ratio was increased to 600 to enable throttling to the low end of the required input-power range. The geometry of the apertures in the grids was selected on the basis of experience in the use of grids of similar geometry in the NSTAR thruster. Characteristics of the aperture geometry include a high open-area fraction in the screen grid to reduce discharge losses and a low open-area fraction in the accelerator grid to reduce losses of electrically neutral gas atoms or molecules. The NEXT accelerator grid was made thicker than that of the NSTAR to make more material available for erosion, thereby increasing the service life and, hence, the total impulse. The NEXT grids are made of molybdenum, which was chosen because its combination of high strength and low thermal expansion helps to minimize thermally and inertially induced

  9. Ion velocity and plasma potential measurements of a cylindrical cusped field thruster

    SciTech Connect

    MacDonald, N. A.; Young, C. V.; Cappelli, M. A.; Hargus, W. A. Jr.

    2012-05-01

    Measurements of the most probable time-averaged axial ion velocities and plasma potential within the acceleration channel and in the plume of a straight-channeled cylindrical cusped field thruster operating on xenon are presented. Ion velocities for the thruster are derived from laser-induced fluorescence measurements of the 5d[4]{sub 7/2}-6p[3]{sub 5/2} xenon ion excited state transition centered at {lambda}=834.72nm. Plasma potential measurements are made using a floating emissive probe with a thoriated-tungsten filament. The thruster is operated in a power matched condition with 300 V applied anode potential for comparison to previous krypton plasma potential measurements, and a low power condition with 150 V applied anode potential. Correlations are seen between the plasma potential drop outside of the thruster and kinetic energy contours of the accelerating ions.

  10. Sputter erosion and deposition in the discharge chamber of a small mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Power, J. L.

    1973-01-01

    A 5-cm diameter mercury ion thruster similar to one tested for 9715 hours was operated approximately 400 hrs each at discharge voltages of 36.6, 39.6, and 42.6 V, with corresponding discharge propellant utilizations of 58, 68, and 70 percent. The observed sputter erosion rates of the internal thruster parts and the anode weight gain rate all rose rapidly with discharge voltage and were roughly in the ratio of 1:3:5 for the three voltages. The combined weight loss of the internal thruster parts nearly balanced the anode weight gain. Hg+2 ions apparently caused most of the observed erosion.

  11. Sputter erosion and deposition in the discharge chamber of a small mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Power, J. L.

    1973-01-01

    A 5 cm diameter mercury ion thruster similar to one tested for 9715 hours was operated approximately 400 hrs each at discharge voltages of 36.6, 39.6, and 42.6 V, with corresponding discharge propellant utilizations of 58, 68, and 70 percent. The observed sputter erosion rates of the internal thruster parts and the anode weight gain rate all rose rapidly with discharge voltage and were roughly in the ratio of 1:3:5 for the three voltages. The combined weight loss of the internal thruster parts nearly balanced the anode weight gain. Hg(+2) ion apparently caused most of the observed erosion.

  12. Inter-cusp Ion and Electron Transport in a Nstar-derivative Ion Thruster

    NASA Technical Reports Server (NTRS)

    Foster, John E.

    2001-01-01

    Diffusion of electrons and ions to anode surfaces between the magnetic cusps of a NASA Solar Electric Propulsion Technology Application Readiness ion thruster has been characterized. Ion flux measurements were made at the anode and at the screen grid electrode. The measurements indicated that the average ion current density at the anode and at the screen grid were approximately equal. Additionally, it was found that the electron flux to the anode between cusps is best described by the classical cross-field diffusion coefficient.

  13. Ion and Electron Transport in an Nstar-derivative Ion Thruster. Revised

    NASA Technical Reports Server (NTRS)

    Foster, John E.

    2001-01-01

    Diffusion of electrons and ions to anode surfaces between the magnetic cusps of a NASA Solar Electric Propulsion Technology Application Readiness ion thruster has been characterized. Ion flux measurements were made at the anode and at the screen grid electrode. The measurements indicated that the average ion current density at the anode and at the screen grid were approximately equal. Additionally, it was found that the electron flux to the anode between cusps is best described by the classical cross-field diffusion coefficient.

  14. Performance capabilities of the 8-cm mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1981-01-01

    A preliminary characterization of the performance capabilities of the 8-cm thruster in order to initiate an evaluation of its application to LSS propulsion requirements is presented. With minor thruster modifications, the thrust was increased by about a factor of four while the discharge voltage was reduced from 39 to 22 volts. The thruster was operated over a range of specific impulse of 1950 to 3040 seconds and a maximum total efficiency of about 54 percent was attained. Preliminary analysis of component lifetimes, as determined by temperature and spectroscopic line intensity measurements, indicated acceptable thruster lifetimes are anticipated at the high power level operation.

  15. Eight cm technology thruster development. [structurally integrated ion thruster for attitude control and stationkeeping of synchronous satellites

    NASA Technical Reports Server (NTRS)

    Hyman, J., Jr.

    1974-01-01

    A structural integrated ion thruster with 8-cm beam diameter (SIT-8) was developed for attitude control and stationkeeping of synchronous satellites. As optimized, the system demonstrates a thrust T=1.14 mlb (not corrected for beam V sub B = 1200 V (I sub sp = 2200 sec) total propellant utilization efficiency nu sub u = 59.8% (is approximately 72% without auxiliary pulse-igniter electrode), and electrical efficiency n sub E 61.9%. The thruster incorporates a wire-mesh anode and tantalum cover surfaces to control discharge chamber flake formation and employs an auxiliary pulse-igniter electrode for hollow-cathode ignition. When the SIT-8 is integrated with the compatible SIT-5 propellant tankage, the system envelope is 35 cm long by 13 cm flange bolt circle with a mass of 9.8 kg including 6.8 kg of mercury propellant. Two thrust vectoring systems which generate beam deflections in two orthogonal directions were also developed under the program and tested with the 8-cm thruster. One system vectors the beam over + or - 10 degrees by gimbaling of the entire thruster (not including tankage), while the other system vectors the beam over + or - 7 degrees by translating the accel electrode relative to the screen electrode.

  16. Production of High Energy Ions Near an Ion Thruster Discharge Hollow Cathode

    NASA Technical Reports Server (NTRS)

    Katz, Ira; Mikellides, I. G.; Goebel, D. M.; Jameson, K. K.; Wirz, R.; Polk, James E.

    2006-01-01

    Several researchers have measured ions leaving ion thruster discharge chambers with energies far greater than measured discharge chamber potentials. Presented in this paper is a new mechanism for the generation of high energy ions and a comparison with measured ion spectra. The source of high energy ions has been a puzzle because they not only have energies in excess of measured steady state potentials, but as reported by Goebel et. al. [1], their flux is independent of the amplitude of time dependent plasma fluctuations. The mechanism relies on the charge exchange neutralization of xenon ions accelerated radially into the potential trough in front of the discharge cathode. Previous researchers [2] have identified the importance of charge exchange in this region as a mechanism for protecting discharge cathode surfaces from ion bombardment. This paper is the first to identify how charge exchange in this region can lead to ion energy enhancement.

  17. Operation of the J-series thruster using inert gas

    NASA Technical Reports Server (NTRS)

    Rawlin, V. K.

    1982-01-01

    Electron bombardment ion thrusters using inert gases are candidates for large space systems. The J-Series 30 cm diameter thruster, designed for operation up to 3 k-W with mercury, is at a state of technology readiness. The characteristics of operation with xenon, krypton, and argon propellants in a J-Series thruster with that obtained with mercury are compared. The performance of the discharge chamber, ion optics, and neutralizer and the overall efficiency as functions of input power and specific impulse and thruster lifetime were evaluated. As expected, the discharge chamber performance with inert gases decreased with decreasing atomic mass. Aspects of the J-Series thruster design which would require modification to provide operation at high power with insert gases were identified.

  18. Measurements of neutral and ion velocity distribution functions in a Hall thruster

    NASA Astrophysics Data System (ADS)

    Svarnas, Panagiotis; Romadanov, Iavn; Diallo, Ahmed; Raitses, Yevgeny

    2015-11-01

    Hall thruster is a plasma device for space propulsion. It utilizes a cross-field discharge to generate a partially ionized weakly collisional plasma with magnetized electrons and non-magnetized ions. The ions are accelerated by the electric field to produce the thrust. There is a relatively large number of studies devoted to characterization of accelerated ions, including measurements of ion velocity distribution function using laser-induced fluorescence diagnostic. Interactions of these accelerated ions with neutral atoms in the thruster and the thruster plume is a subject of on-going studies, which require combined monitoring of ion and neutral velocity distributions. Herein, laser-induced fluorescence technique has been employed to study neutral and single-charged ion velocity distribution functions in a 200 W cylindrical Hall thruster operating with xenon propellant. An optical system is installed in the vacuum chamber enabling spatially resolved axial velocity measurements. The fluorescence signals are well separated from the plasma background emission by modulating the laser beam and using lock-in detectors. Measured velocity distribution functions of neutral atoms and ions at different operating parameters of the thruster are reported and analyzed. This work was supported by DOE contract DE-AC02-09CH11466.

  19. The use of electrostatic probes to characterize the discharge plasma structure and identify discharge cathode erosion mechanisms in ring-cusp ion thrusters

    NASA Astrophysics Data System (ADS)

    Herman, Daniel Andrew

    The erosion of the discharge cathode assembly (DCA) is currently one of the lifetime limiting factors of ion thruster operation and will play an even more important role for more ambitious, future ion thruster applications requiring more throughput at higher-power. Erosion of the DCA has been observed throughout the ground-based wear testing of the 30-cm NSTAR ion thruster. Energetic ions have been detected near the DCA, from Laser-Induced Fluorescence (LIF) measurements, that appear to be the cause of the DCA erosion, though a mechanism by which ions gain sufficient energy to sputter erode the DCA material has not been determined. This dissertation presents research aimed at characterizing the discharge chamber plasma near the DCA to determine the mechanism by which energetic ions are created and erode the DCA inside ring-cusp ion engines. A diagnostic technique is developed to interrogate the near-DCA regions of two ion thrusters: the 30-cm FMT2 NSTAR and 40-cm LM4 NEXT engines. Both engines contain similar plasma structures. Number densities are highest along cathode centerline as the axial magnetic field near the DCA effectively confines electrons to a narrow plume. Plasma potential mappings rule out the existence of a potential-hill that has been proposed as the cause of the DCA erosion. A free standing potential gradient structure is found to form the transition between the low-potential cathode plume and the high-potential bulk discharge plasma, termed a double layer. The field-aligned double layer accelerates ions from the bulk discharge plasma towards the DCA centerline. Measured plasma parameters and LIF velocimetry data are used to calculate an erosion rate utilizing near-threshold sputtering yield formulae. Singly-ionized xenon cannot solely account for the observed NSTAR erosion rates. Incorporation of double-ionized xenon from measured double-to-single current measurements in the plume of the 30-cm and 40-cm thrusters significantly increases the

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

    NASA Technical Reports Server (NTRS)

    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.

  1. Ion propulsion cost effectivity

    NASA Technical Reports Server (NTRS)

    Zafran, S.; Biess, J. J.

    1978-01-01

    Ion propulsion modules employing 8-cm thrusters and 30-cm thrusters were studied for Multimission Modular Spacecraft (MMS) applications. Recurring and nonrecurring cost elements were generated for these modules. As a result, ion propulsion cost drivers were identified to be Shuttle charges, solar array, power processing, and thruster costs. Cost effective design approaches included short length module configurations, array power sharing, operation at reduced thruster input power, simplified power processing units, and power processor output switching. The MMS mission model employed indicated that nonrecurring costs have to be shared with other programs unless the mission model grows. Extended performance missions exhibited the greatest benefits when compared with monopropellant hydrazine propulsion.

  2. Electron Transport and Ion Acceleration in a Low-power Cylindrical Hall Thruster

    SciTech Connect

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

    2004-06-24

    Conventional annular Hall thrusters become inefficient when scaled to low power. Cylindrical Hall thrusters, which have lower surface-to-volume ratio, are therefore more promising for scaling down. They presently exhibit performance comparable with conventional annular Hall thrusters. Electron cross-field transport in a 2.6 cm miniaturized cylindrical Hall thruster (100 W power level) has been studied through the analysis of experimental data and Monte Carlo simulations of electron dynamics in the thruster channel. The numerical model takes into account elastic and inelastic electron collisions with atoms, electron-wall collisions, including secondary electron emission, and Bohm diffusion. We show that in order to explain the observed discharge current, the electron anomalous collision frequency {nu}{sub B} has to be on the order of the Bohm value, {nu}{sub B} {approx} {omega}{sub c}/16. The contribution of electron-wall collisions to cross-field transport is found to be insignificant. The plasma density peak observed at the axis of the 2.6 cm cylindrical Hall thruster is likely to be due to the convergent flux of ions, which are born in the annular part of the channel and accelerated towards the thruster axis.

  3. Numerical Modeling of Ion Dynamics in a Carbon Nanotube Field-Ionized Thruster

    DTIC Science & Technology

    2011-12-01

    ION SOURCES Ions may be produced by several methods: photo-ionization, electron bombardment, field ionization, surface ionization, and thermionic ...OF ION DYNAMICS IN A CARBON NANOTUBE FIELD -IONIZED ION THRUSTER by Sarah F. Michael December 2011 Thesis Advisors: Dragoslav Grbovic...December 2011 3. REPORT TYPE AND DATES COVERED Master’s Thesis 4. TITLE AND SUBTITLE Numerical Modeling of Ion Dynamics in a Carbon Nanotube Field

  4. Specific spacecraft evaluation: Special report. [charged particle transport from a mercury ion thruster to spacecraft surfaces

    NASA Technical Reports Server (NTRS)

    Sellen, J. M., Jr.

    1978-01-01

    Charged and neutral particle transport from an 8 cm mercury ion thruster to the surfaces of the P 80-1 spacecraft and to the Teal Ruby sensor and the ECOM-501 sensor of that spacecraft were investigated. Laboratory measurements and analyses were used to examine line-of-sight and nonline-of sight particle transport modes. The recirculation of Hg(+) ions in the magnetic field of the earth was analyzed for spacecraft velocity and Earth magnetic field vector configurations which are expected to occur in near Earth, circular, high inclination orbits. For these magnetic field and orbit conditions and for expected ion release distribution functions, in both angles and energies, the recirculation/re-interception of ions on spacecraft surfaces was evaluated. The refraction of weakly energetic ions in the electric fields of the thruster plasma plume and in the electric fields between this plasma plume and the material boundaries of the thruster, the thruster sputter shield, and the various spacecraft surfaces were examined. The neutral particle transport modes of interest were identified as sputtered metal atoms from the thruster beam shield. Results, conclusions, and future considerations are presented.

  5. Mission Benefits of Gridded Ion and Hall Thruster Hybrid Propulsion Systems

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.; Polsgrove, Tara

    2006-01-01

    The NASA In-Space Propulsion Technology (ISPT) Project Office has been developing the NEXT gridded ion thruster system and is planning to procure a low power Hall system. The new ion propulsion systems will join NSTAR as NASA's primary electric propulsion system options. Studies have been performed to show mission benefits of each of the stand alone systems. A hybrid ion propulsion system (IPS) can have the advantage of reduced cost, decreased flight time and greater science payload delivery over comparable homogeneous systems. This paper explores possible advantages of combining various thruster options for a single mission.

  6. Global model of a gridded-ion thruster powered by a radiofrequency inductive coil

    SciTech Connect

    Chabert, P.; Arancibia Monreal, J.; Bredin, J.; Popelier, L.; Aanesland, A.

    2012-07-15

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

  7. Transport of Sputtered Carbon During Ground-Based Life Testing of Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Marker, Colin L.; Clemons, Lucas A.; Banks, Bruce A.; Miller, Sharon; Snyder, Aaron; Hung, Ching-Cheh; Karniotis, Christina A.; Waters, Deborah L.

    2005-01-01

    High voltage, high power electron bombardment ion thrusters needed for deep space missions will be required to be operated for long durations in space as well as during ground laboratory life testing. Carbon based ion optics are being considered for such thrusters. The sputter deposition of carbon and arc vaporized carbon flakes from long duration operation of ion thrusters can result in deposition on insulating surfaces, causing them to become conducting. Because the sticking coefficient is less than one, secondary deposition needs to be considered to assure that shorting of critical components does not occur. The sticking coefficient for sputtered carbon and arc vaporized carbon is measured as well as directional ejection distribution data for carbon that does not stick upon first impact.

  8. Increasing Extracted Beam Current Density in Ion Thrusters through Plasma Potential Modification

    NASA Astrophysics Data System (ADS)

    Arthur, Neil; Foster, John

    2015-09-01

    A gridded ion thruster's maximum extractable beam current is determined by the space charge limit. The classical formulation does not take into account finite ion drift into the acceleration gap. It can be shown that extractable beam current can be increased beyond the conventional Child-Langmuir law if the ions enter the gap at a finite drift speed. In this work, ion drift in a 10 cm thruster is varied by adjusting the plasma potential relative to the potential at the extraction plane. Internal plasma potential variations are achieved using a novel approach involving biasing the magnetic cusps. Ion flow variations are assessed using simulated beam extraction in conjunction with a retarding potential analyzer. Ion beam current density changes at a given total beam voltage in full beam extraction tests are characterized as a function of induced ion drift velocity as well.

  9. A coupled performance and thermal model for radio-frequency gridded ion thrusters*

    NASA Astrophysics Data System (ADS)

    Dobkevicius, Mantas; Feili, Davar

    2016-10-01

    Recently proposed space missions such as Darwin, eLISA and NGGM have encouraged the development of electric propulsion thrusters capable of operating in the micro-Newton (μN) thrust range. To meet these requirements, radio frequency (RF) gridded-ion thrusters need to be scaled down to a few centimeters in size. Due to the small size of these thrusters, it is important to accurately determine the thermal and performance parameters. To achieve this, a multi-physics performance model has been developed, composed of plasma discharge, 2D axisymmetric ion extraction, 3D electromagnetic and RF circuit models. The plasma discharge model itself is represented using 0D global, 2D axisymmetric and 3D molecular neutral gas, and Boltzmann electron transport sub-models. A 3D thermal model is introduced to determine the temperature distribution for various throttle points, using as inputs the plasma and electromagnetic field heating values obtained from the performance model. This also allows the validation of the performance model itself. Additionally, we analyze the effect the thruster's temperatures play on the plasma properties/performance and vice versa. The model is based on the RIT 3.5 thruster developed for the NGGM mission geometry and predicts the RIT 3.5 experimental data within approximately 10%. Contribution to the Topical Issue "Physics of Ion Beam Sources", edited by Holger Kersten and Horst Neumann.

  10. Ion optics for high power 50-cm-diam ion thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.; Millis, Marc G.

    1989-01-01

    The process used at the NASA-Lewis to fabricate 30 and 50-cm-diameter ion optics is described. The ion extraction capabilities of the 30 and 50-cm diameter ion optics were evaluated on divergent field and ring-cusp discharge chambers and compared. Perveance was found to be sensitive to the effects of the type and power of the discharge chamber and to the accelerator electrode hole diameter. Levels of up to 0.64 N and 20 kW for thrust and input power, respectively, were demonstrated with the divergent-field discharge chamber. Thruster efficiencies and specific impulse values up to 79 percent and 5000 sec., respectively, were achieved with the ring-cusp discharge chamber.

  11. Ion ejection from a permanent-magnet mini-helicon thruster

    SciTech Connect

    Chen, Francis F.

    2014-09-15

    A small helicon source, 5 cm in diameter and 5 cm long, using a permanent magnet (PM) to create the DC magnetic field B, is investigated for its possible use as an ion spacecraft thruster. Such ambipolar thrusters do not require a separate electron source for neutralization. The discharge is placed in the far-field of the annular PM, where B is fairly uniform. The plasma is ejected into a large chamber, where the ion energy distribution is measured with a retarding-field energy analyzer. The resulting specific impulse is lower than that of Hall thrusters but can easily be increased to relevant values by applying to the endplate of the discharge a small voltage relative to spacecraft ground.

  12. Charge-exchange erosion studies of accelerator grids in ion thrusters

    NASA Technical Reports Server (NTRS)

    Peng, Xiaohang; Ruyten, Wilhelmus M.; Keefer, Dennis

    1993-01-01

    A particle simulation model is developed to study the charge-exchange grid erosion in ion thrusters for both ground-based and space-based operations. Because the neutral gas downstream from the accelerator grid is different for space and ground operation conditions, the charge-exchange erosion processes are also different. Based on an assumption of now electric potential hill downstream from the ion thruster, the calculations show that the accelerator grid erosion rate for space-based operating conditions should be significantly less than experimentally observed erosion rates from the ground-based tests conducted at NASA Lewis Research Center (LeRC) and NASA Jet Propulsion Laboratory (JPL). To resolve this erosion issue completely, we believe that it is necessary to accurately measure the entire electric potential field downstream from the thruster.

  13. Structural Analysis of Pyrolytic Graphite Optics for the HiPEP Ion Thruster

    NASA Technical Reports Server (NTRS)

    Meckel, Nicole; Polaha, Jonathan; Juhlin, Nils

    2006-01-01

    The long lifetime requirements of interplanetary exploration missions is driving the need to develop long-life components for the electric propulsion thrusters that are being targeted for these missions. One of the primary life-limiting components of ion thrusters are the optics, which are continuously eroded during the operation of the thruster. Pyrolytic graphite optics are being considered for the High Power Electric Propulsion (HiPEP) ion thruster because of their very high resistance to erosion. This paper describes the structural analysis of the HiPEP pyrolytic graphite. A description of the development of the grid model, as well as the development of the effective properties and stress concentrations in the apertured area of the grids is included. An evaluation of the use of curved grids shows that the increased stiffness (compared to flat grids) prevents intergrid impact during launch, however, the residual stresses introduced by curving the grids pushes the resulting peak stresses beyond the critical stress. As a result, flat grids are recommended as the design solution. Thermally induced grid displacements during normal thruster operation are also presented.

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

    NASA Astrophysics Data System (ADS)

    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

  15. Multi-Thruster Propulsion Apparatus

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J. (Inventor)

    2016-01-01

    An electric propulsion machine includes an ion thruster having a discharge chamber housing a large surface area anode. The ion thruster includes flat annular ion optics with a small span to gap ratio. Optionally, at least a second thruster may be disposed radially offset from the ion thruster.

  16. Experimental investigation of discharge plasma magnetic confinement in the NSTASR ion thruster

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Fitzgerald, Dennis; Owens, Al

    2005-01-01

    Magnetic confinement studies were performed on the state-of-the-art NSTAR ion thruster. The goal of the experimental studies was determine the dependence of plasma confinement and plasma uniformity on the strength and shape of the imposed ring-cusp magnetic field.

  17. Status of the extended life test of the Deep Space 1 flight spare ion engine after 30,352 hours of operation

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Brophy, John R.; Goodfellow, Keith D.

    2003-01-01

    The extended life test (ELT) of the Deep Space 1 (DS1) spare flight ion thruster (FT2) was voluntarily terminated June 26th 2003. The test was started in Cotober of 1998, just prior to the launch of the DS1 spacecraft, with the primary purpose of determining the ultimate service life capability of the NASA 30-cm-ion thruster technology.

  18. Evaluation of ion current density distribution on an extraction electrode of a radio frequency ion thruster

    NASA Astrophysics Data System (ADS)

    Masherov, P.; Riaby, V.; Abgaryan, V.

    2017-01-01

    The radial distributions of ion current density on an ion extracting electrode of a radio frequency (RF) ion thruster (RIT) with an inductive plasma source were obtained using probe diagnostics of the RF xenon plasma. Measurements were carried out using a plane wall probe simulator and the VGPS-12 Probe System of Plasma Sensors Co. At xenon flow rate q  =  2 sccm plasma pressure was 2 · 10-3 Torr, incident RF generator power varied in the range P g  =  50-250 W with RF power absorbed by plasma up to P p  =  220 W. Ion current densities were determined using semi- and double-logarithmic probe characteristics by linear extrapolations of their ion branches to probe floating potentials. The same parameters were also measured in undisturbed plasma by a classic cylindrical probe. They exceeded plane probe data by more than two times, showing the effectiveness of plasma sheath reproduction of the RIT ion extracting electrode by the plane wall probe simulator. Slight non-uniformity of the resulting plasma distributions and simplified RIT model design showed that the studied device with flat antenna coil and ferrite core could be considered as a promising prospect for RITs of new generation.

  19. Visual evidence of suppressing the ion and electron energy loss on the wall in Hall thrusters

    NASA Astrophysics Data System (ADS)

    Ding, Yongjie; Peng, Wuji; Sun, Hezhi; Wei, Liqiu; Zeng, Ming; Wang, Fufeng; Yu, Daren

    2017-03-01

    A method of pushing down magnetic field with two permanent magnetic rings is proposed in this paper. It can realize ionization in a channel and acceleration outside the channel. The wall will only suffer from the bombardment of low-energy ions and electrons, which can effectively reduce channel erosion and extend the operational lifetime of thrusters. Furthermore, there is no additional power consumption of coils, which improves the efficiency of systems. We show here the newly developed 200 W no wall-loss Hall thruster (NWLHT-200) that applies the method of pushing down magnetic field with two permanent magnetic rings; the visual evidence we obtained preliminarily confirms the feasibility that the proposed method can realize discharge without wall energy loss or erosion of Hall thrusters.

  20. Particle and field measurements on two J-series 30-centimeter thrusters

    NASA Technical Reports Server (NTRS)

    Lathem, W. C.

    1981-01-01

    An experimental investigation under complete computer control of the characteristics of particles and fields emanating from mercury bombardment ion thrusters (J-series, 30 cm diameter), operating independently and simultaneously, is discussed. Results show that the flux rate of sputtered metal atoms leaving a single thruster has been determined and the data is in agreement with previous test results. The flux rate and energies of ions leaving single and two-thruster configurations have been established for a variety of beam current levels and combinations involving the two thrusters, and the data are sufficient to establish upper bounds. Floating potential measurements made in the near vicinity of the thrusters indicate that the fields are weak, plus or minus a few volts. The tests may be significant for the Solar Electric Propulsion System (SEPS) planners.

  1. Optical Studies Accompanying a Performance Test of a HEMP Thruster in the JUMBO Test Facility

    NASA Astrophysics Data System (ADS)

    Weis, S.; Meiß, S.; Schartner, K.-H.; Feili, D.; Harmann, H.-P.; Koch, N.; Kornfeld, G.

    2004-10-01

    Ion thrusters are bright light sources allowing a spectral analysis in a short time without disturbance of the plume. Accordingly, the light emission of a Thales HEMP3) thruster DM6 was studied while the performance of the thruster was investigated in the JUMBO fest facility of the Giessen University. The front end of a light waveguide was positioned in the demagnified image of the thruster exit diaphragm and its other end fixed in the entrance slit of a monochromator of 30 cm focal length, equipped with a CCD as detector. The HEMP thruster was mounted on a thrust balance. The angular distribution of the emitted ions could be monitored by a three axis scanner. Mass flow and discharge voltage were varied during the performance test. The collected spectra contain lines from neutral, singly and doubly charged Xe serving as propellant. The dependence of the relative intensities of selected lines on the discharge parameters will be discussed.

  2. NASA's Evolutionary Xenon Thruster: The NEXT Ion Propulsion System for Solar System Exploration

    NASA Technical Reports Server (NTRS)

    Pencil, Eric J.; Benson, Scott W.

    2008-01-01

    This viewgraph presentation reviews NASA s Evolutionary Xenon Thruster (NEXT) Ion Propulsion system. The NEXT project is developing a solar electric ion propulsion system. The NEXT project is advancing the capability of ion propulsion to meet NASA robotic science mission needs. The NEXT system is planned to significantly improve performance over the state of the art electric propulsion systems, such as NASA Solar Electric Propulsion Technology Application Readiness (NSTAR). The status of NEXT development is reviewed, including information on the NEXT Thruster, the power processing unit, the propellant management system (PMS), the digital control interface unit, and the gimbal. Block diagrams NEXT system are presented. Also a review of the lessons learned from the Dawn and NSTAR systems is provided. In summary the NEXT project activities through 2007 have brought next-generation ion propulsion technology to a sufficient maturity level.

  3. Plume and Discharge Plasma Measurements of an NSTAR-type Ion Thruster

    NASA Technical Reports Server (NTRS)

    Foster, John E; Soulas, George C.; Patterson, Michael J.

    2000-01-01

    The success of the NASA Deep Space I spacecraft has demonstrated that ion propulsion is a viable option for deep space science missions. More aggressive missions such as Comet Nuclear Sample Return and Europa lander will require higher power, higher propellant throughput and longer thruster lifetime than the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) engine. Presented here are thruster plume and discharge plasma measurements of an NSTAR-type thruster operated from 0.5 kW to 5 kW. From Faraday plume sweeps, beam divergence was determined. From Langmuir probe plume measurements on centerline, low energy ion production on axis due to charge-exchange and direct ionization was assessed. Additionally, plume plasma potential measurements made on axis were used to determine the upper energy limits at which ions created on centerline could be radially accelerated. Wall probes flush-mounted to the thruster discharge chamber anode were used to assess plasma conditions. Langmuir probe measurements at the wall indicated significant differences in the electron temperature in the cylindrical and conical sections of the discharge chamber.

  4. Ion Current Density Study of the NASA-300M and NASA-457Mv2 Hall Thrusters

    NASA Technical Reports Server (NTRS)

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

    2012-01-01

    NASA Glenn Research Center is developing a Hall thruster in the 15-50 kW range to support future NASA missions. As a part of the process, the performance and plume characteristics of the NASA-300M, a 20-kW Hall thruster, and the NASA-457Mv2, a 50-kW Hall thruster, were evaluated. The collected data will be used to improve the fidelity of the JPL modeling tool, Hall2De, which will then be used to aid the design of the 15-50 kW Hall thruster. This paper gives a detailed overview of the Faraday probe portion of the plume characterization study. The Faraday probe in this study is a near-field probe swept radially at many axial locations downstream of the thruster exit plane. Threshold-based integration limits with threshold values of 1/e, 1/e2, and 1/e3 times the local peak current density are tried for the purpose of ion current integration and divergence angle calculation. The NASA-300M is operated at 7 conditions and the NASA-457Mv2 at 14 conditions. These conditions span discharge voltages of 200 to 500 V and discharge power of 10 to 50 kW. The ion current density profiles of the near-field plume originating from the discharge channel are discovered to strongly resemble Gaussian distributions. A novel analysis approach involving a form of ray tracing is used to determine an effective point of origin for the near-field plume. In the process of performing this analysis, definitive evidence is discovered that showed the near-field plume is bending towards the thruster centerline.

  5. Ion Current Density Study of the NASA-300M and NASA-457Mv2 Hall Thrusters

    NASA Technical Reports Server (NTRS)

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

    2012-01-01

    NASA Glenn Research Center is developing a Hall thruster in the 15-50 kW range to support future NASA missions. As a part of the process, the performance and plume characteristics of the NASA-300M, a 20-kW Hall thruster, and the NASA-457Mv2, a 50-kW Hall thruster, were evaluated. The collected data will be used to improve the fidelity of the JPL modeling tool, Hall2De, which will then be used to aid the design of the 15-50 kW Hall thruster. This paper gives a detailed overview of the Faraday probe portion of the plume characterization study. The Faraday probe in this study is a near-field probe swept radially at many axial locations downstream of the thruster exit plane. Threshold-based integration limits with threshold values of 1/e, 1/e(sup 2), and 1/e(sup 3) times the local peak current density are tried for the purpose of ion current integration and divergence angle calculation. The NASA-300M is operated at 7 conditions and the NASA-457Mv2 at 14 conditions. These conditions span discharge voltages of 200 to 500 V and discharge power of 10 to 50 kW. The ion current density profiles of the near-field plume originating from the discharge channel are discovered to strongly resemble Gaussian distributions. A novel analysis approach involving a form of ray tracing is used to determine an effective point of origin for the near-field plume. In the process of performing this analysis, definitive evidence is discovered that showed the near-field plume is bending towards the thruster centerline.

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

    NASA Technical Reports Server (NTRS)

    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.

  7. Diagnostic of plasma streams from ion thrusters for space propulsion using emissive probes

    NASA Astrophysics Data System (ADS)

    Conde, L.; Tierno, S. P.; Domenech-Garret, J. L.; Donoso, J. M.; Castillo, M. A.; Eíriz, I.; Sáez de Ocáriz, I.

    2016-10-01

    The emissive probes are employed for the determination of the local plasma potential of plasma streams produced by ion thrusters. Its operation basically relies on electron collection and emission and are less sensitive to the ion motion than collecting probes. The diagnostic using emissive probes is reviewed with emphasis in low density plasmas. Our results support the conclusion that potential structures around the probe, as virtual cathodes, would be responsible for the operation of emissive probes in low density plasmas.

  8. Inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1976-01-01

    Inert gases are of interest as possible alternatives to the usual electric thruster propellants of mercury and cesium. The multipole discharge chamber investigated was shown capable of low discharge chamber losses and flat ion beam profiles with a minimum of optimization. Minimum discharge losses were 200 to 250 eV/ion for xenon and 300 to 350 eV/ion for argon, while flatness parameters in the plane of the accelerator grid were 0.85 to 0.95. The design used employs low magnetic field strengths, which permits the use of sheet-metal parts. The corner problem of the discharge chamber was resolved with recessed corner anodes, which approximately equalized both the magnetic field above the anodes and the electron currents to these anodes. Argon hollow cathodes were investigated at currents up to about 5 amperes using internal thermionic emitters. Cathode chamber diameter optimized in the 1.0 to 2.5 cm range, while orifices diameter optimized in the 0.5 to 5 mm range. The use of a bias voltage for the internal emitter extended the operating range and facilitated starting. The masses of 15 and 30 cm flight type thrusters were estimated at about 4.2 and 10.8 kg.

  9. The UK Ion Thruster System and a Proposed Future Programme.

    DTIC Science & Technology

    1977-05-01

    As is norma l in the development of a comp lex sys tem, certain components must be defined almost comp le tel y before it is sensible to proceed far...outputs from T4A thruster power conditioning uni t Parameter Accuracy Norma l value Operating range Beam supply voltage ±2% 900V Stabi li sed 2 Beam...experienced under norma l operating conditions , and the possibili ty of its occurrence is considered small; the factor 4 safety marg in on minimum

  10. Determination of the force transmitted by an ion thruster plasma plume to an orbital object

    NASA Astrophysics Data System (ADS)

    Alpatov, A.; Cichocki, F.; Fokov, A.; Khoroshylov, S.; Merino, M.; Zakrzhevskii, A.

    2016-02-01

    An approach to determine the force transmitted by the plasma plume of an ion thruster to an orbital object immersed in it using its central projection on a selected plane is proposed. A photo camera is used to obtain the image of the object central projection. The algorithms for the calculation of the transmission of momentum by the impacting ion beam are developed including the determination of the object contour and the correction of the error due to a camera offset from the ion beam axis, and the computation of the fraction of the ion beam that impinges on the object surface.

  11. Modeling of surface-dominated plasmas: from electric thruster to negative ion source.

    PubMed

    Taccogna, F; Schneider, R; Longo, S; Capitelli, M

    2008-02-01

    This contribution shows two important applications of the particle-in-cell/monte Carlo technique on ion sources: modeling of the Hall thruster SPT-100 for space propulsion and of the rf negative ion source for ITER neutral beam injection. In the first case translational degrees of freedom are involved, while in the second case inner degrees of freedom (vibrational levels) are excited. Computational results show how in both cases, plasma-wall and gas-wall interactions play a dominant role. These are secondary electron emission from the lateral ceramic wall of SPT-100 and electron capture from caesiated surfaces by positive ions and atoms in the rf negative ion source.

  12. An advanced electric propulsion diagnostic (AEPD) platform for in-situ characterization of electric propulsion thrusters and ion beam sources

    NASA Astrophysics Data System (ADS)

    Bundesmann, Carsten; Eichhorn, Christoph; Scholze, Frank; Spemann, Daniel; Neumann, Horst; Pagano, Damiano; Scaranzin, Simone; Scortecci, Fabrizio; Leiter, Hans J.; Gauter, Sven; Wiese, Ruben; Kersten, Holger; Holste, Kristof; Köhler, Peter; Klar, Peter J.; Mazouffre, Stéphane; Blott, Richard; Bulit, Alexandra; Dannenmayer, Käthe

    2016-10-01

    Experimental characterization is an essential task in development, qualification and optimization process of electric propulsion thrusters or ion beam sources for material processing, because it can verify that the thruster or ion beam source fulfills the requested mission or application requirements, and it can provide parameters for thruster and plasma modeling. Moreover, there is a need for standardizing electric propulsion thruster diagnostics in order to make characterization results of different thrusters and also from measurements performed in different vacuum facilities reliable and comparable. Therefore, we have developed an advanced electric propulsion diagnostic (AEPD) platform, which allows a comprehensive in-situ characterization of electric propulsion thrusters (or ion beam sources) and could serve as a standard on-ground tool in the future. The AEPD platform uses a five-axis positioning system and provides the option to use diagnostic tools for beam characterization (Faraday probe, retarding potential analyzer, ExB probe, active thermal probe), for optical inspection (telemicroscope, triangular laser head), and for thermal characterization (pyrometer, thermocamera). Here we describe the capabilities of the diagnostic platform and provide first experimental results of the characterization of a gridded ion thruster RIT- μX.

  13. The 2.5 kW advanced technology ion thruster. [design analysis and performance tests

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.

    1976-01-01

    A representative thruster was extensively documented with respect to performance parameters and characteristics at selected ion beam currents in the 0.5 to 2.75 A range, including measurements of thrust losses resulting from doubly-charged ions and ion beam divergence. Corrected total efficiency was shown to be relatively insensitive to operating parameter selection at any given power level. Factors affecting doubly-charged ionization were studied and it was found that the fraction of doubly-charged ions is directly proportional to the discharge chamber propellant utilization. The parameter that most affects this proportionality is the accel aperture diameter (which controls neutral atom loss). Thruster-power conditioner interactions were studied with the result that previous power supply specifications remain satisfactory. Options for reducing the number of power supplies required were demonstrated to be feasible. Gimbal actuator designs were studied with the goal of selecting a particular approach for design and development. The conclusion drawn was that optimum gimbal actuator design depends heavily on the thruster application and consequently the effort was concluded by developing a computer program to aid in specifying the gimbal requirements for the thrust vectoring required in a specific application.

  14. Recent Investigations, Development and Industrial Applications of RF-ion Thrusters in Germany

    NASA Astrophysics Data System (ADS)

    Bassner, H.; Killinger, R.; Kukies, R.; Mueller, J.

    2002-01-01

    R &D work on ion thrusters using radio frequency propellant ionisation has been done at Giessen University since 1962. Engines with ionizer diameters from 4 cm to 35 cm have been designed, built, and tested. Plasma and beam diagnostics have been done and several application studies were carried out. The present work at the 1. Institute of Physics is mainly focussed on two topics: First: The RF-plasma is being modelled in detail in order to establish scaling laws which will allow to scale the existing hardware and save D &Q work. Determination of basic plasma effects shall allow to find the optimum geometry and working parameters (e.g. the discharge vessel length, the best rf-frequency and the necessary discharge pressure). Then, reliable thrust, power, flow rate and efficiency data of different sized RIT-systems can be predicted. Second: The large Giessen test facility "P100,000" (with 30 m3 of chamber volume) has been completely refurbished, to allow the operation of Astriums new 22 cm diam thruster RIT-XT: The oil diffusion pumps were replaced by several cryopumps and two additional turbomolecular pumps. The conical beam target (stainless steel) has been replaced by inclined carbon collector strips. A new beam scanning system has been installed, which can be moved during operation in y- and z- direction. The scanner consists of 160 specially designed Faraday cups. Other diagnostic elements like calorimetric systems, small mass spectrometers etc. can be installed on the scanner, too. The development in the industry did start at Astrium (former MBB) in 1970 with the investigation of the 10 cm discharge chamber diameter laboratory thruster (RIT 10) manufactured in Giessen, using Mercury as the propellant. A first vibration test and a 1000-h lifetime test was performed and did show that the thruster can be used in space. The first 10 years of industrial development were filled with development of electronics, propellant feed system, thruster design and additional

  15. Ion Species Fractions in the Far-Field Plume of a High-Specific Impulse Hall Thruster

    NASA Technical Reports Server (NTRS)

    Hofer, Richard R.; Gallimore, Alec D.

    2003-01-01

    An ExB probe was used to measure the ion species fractions of Xe(+), Xe(2+), and Xe(3+) in the far-field plume of the NASA-173Mv2 laboratory-model Hall thruster. The thruster was operated at a constant xenon flow rate of 10 milligrams per second and discharge voltages of 300 to 900 V. The ExB probe was placed two meters downstream of the thruster exit plane on the thruster centerline. At a discharge voltage of 300 V, the species fractions of Xe(2+) and Xe(3+) were lower, but still consistent with, previous Hall thruster studies using other mass analyzers. Over discharge voltages of 300 to 900 V, the Xe(2+) species fractions increased from 0.04 to 0.12 and the Xe(3+) species fraction increased from 0.01 to 0.02.

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

    NASA Technical Reports Server (NTRS)

    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.

  17. Accelerated life test of sputtering and anode deposit spalling in a small mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Power, J. L.

    1975-01-01

    Tantalum and molybdenum sputtered from discharge chamber components during operation of a 5 centimeter diameter mercury ion thruster adhered much more strongly to coarsely grit blasted anode surfaces than to standard surfaces. Spalling of the sputtered coating did occur from a coarse screen anode surface but only in flakes less than a mesh unit long. The results were obtained in a 200 hour accelerated life test conducted at an elevated discharge potential of 64.6 volts. The test approximately reproduced the major sputter erosion and deposition effects that occur under normal operation but at approximately 75 times the normal rate. No discharge chamber component suffered sufficient erosion in the test to threaten its structural integrity or further serviceability. The test indicated that the use of tantalum-surfaced discharge chamber components in conjunction with a fine wire screen anode surface should cure the problems of sputter erosion and sputtered deposits spalling in long term operation of small mercury ion thrusters.

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

    NASA Technical Reports Server (NTRS)

    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.

  19. Additional application of the NASCAP code. Volume 2: SEPS, ion thruster neutralization and electrostatic antenna model

    NASA Technical Reports Server (NTRS)

    Katz, I.; Cassidy, J. J.; Mandell, M. J.; Parks, D. E.; Schnuelle, G. W.; Stannard, P. R.; Steen, P. G.

    1981-01-01

    The interactions of spacecraft systems with the surrounding plasma environment were studied analytically for three cases of current interest: calculating the impact of spacecraft generated plasmas on the main power system of a baseline solar electric propulsion stage (SEPS), modeling the physics of the neutralization of an ion thruster beam by a plasma bridge, and examining the physical and electrical effects of orbital ambient plasmas on the operation of an electrostatically controlled membrane mirror. In order to perform these studies, the NASA charging analyzer program (NASCAP) was used as well as several other computer models and analytical estimates. The main result of the SEPS study was to show how charge exchange ion expansion can create a conducting channel between the thrusters and the solar arrays. A fluid-like model was able to predict plasma potentials and temperatures measured near the main beam of an ion thruster and in the vicinity of a hollow cathode neutralizer. Power losses due to plasma currents were shown to be substantial for several proposed electrostatic antenna designs.

  20. Ultra High Voltage Propellant Isolators and Insulators for JIMO Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Banks, Bruce A.; Gaier, James R.; Hung, Ching-Cheh; Walters, Patty A.; Sechkar, Ed; Panko, Scott; Kamiotis, Christina A.

    2004-01-01

    Within NASA's Project Prometheus, high specific impulse ion thrusters for electric propulsion of spacecraft for the proposed Jupiter Icy Moon Orbiter (JIMO) mission to three of Jupiter's moons: Callisto, Ganymede and Europa will require high voltage operation to meet mission propulsion. The anticipated approx.6,500 volt net ion energy will require electrical insulation and propellant isolation which must exceed that used successfully by the NASA Solar Electric Propulsion Technology Readiness (NSTAR) Deep Space 1 mission thruster by a factor of approx.6. Xenon propellant isolator prototypes that operate at near one atmosphere and prototypes that operate at low pressures (<100 Torr) have been designed and are being tested for suitability to the JIMO mission requirements. Propellant isolators must be durable to Paschen breakdown, sputter contamination, high temperature, and high voltage while operating for factors longer duration than for the Deep Space 1 Mission. Insulators used to mount the thrusters as well as those needed to support the ion optics have also been designed and are under evaluation. Isolator and insulator concepts, design issues, design guidelines, fabrication considerations and performance issues are presented. The objective of the investigation was to identify candidate isolators and insulators that are sufficiently robust to perform durably and reliably during the proposed JIMO mission.

  1. Simulated Beam Extraction Performance Characterization of a 50-cm Ion Thruster Discharge

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Hubble, Aimee; Nowak-Gucker, Sarah; Davis, Chris; Peterson, Peter; Viges, Eric; Chen, Dave

    2013-01-01

    A 50 cm ion thruster is being developed to operate at >65 percent total efficiency at 11 kW, 2700 s Isp and over 25 kW, 4500 s Isp at a total efficiency of >75 percent. The engine is being developed to address the need for a multimode system that can provide a range of thrust-to- power to service national and commercial near-earth onboard propulsion needs such as station-keeping and orbit transfer. Operating characteristics of the 50 cm ion thruster were measured under simulated beam extraction. The discharge current distribution at the various magnet rings was measured over a range of operating conditions. The relationship between the anode current distribution and the resulting plasma uniformity and ion flux measured at the thruster exit plane is discussed. The thermal envelope will also be investigated through the monitoring of magnet temperatures over the range of discharge powers investigated. Discharge losses as a function of propellant utilization was also characterized at multiple simulated beam currents. Bulk plasma conditions such as electron temperature and electron density near engine centerline was measured over a range of operating conditions using an internal Langmuir probe. Sensitivity of discharge performance to chamber length is also discussed. This data acquired from this discharge study will be used in the refinement of a throttle table in anticipation for eventual beam extraction testing.

  2. Assessment of Spectroscopic, Real-time Ion Thruster Grid Erosion-rate Measurements

    NASA Technical Reports Server (NTRS)

    Domonkos, Matthew T.; Stevens, Richard E.

    2000-01-01

    The success of the ion thruster on the Deep Space One mission has opened the gate to the use of primary ion propulsion. Many of the projected planetary missions require throughput and specific impulse beyond those qualified to date. Spectroscopic, real-time ion thruster grid erosion-rate measurements are currently in development at the NASA Glenn Research Center. A preliminary investigation of the emission spectra from an NSTAR derivative thruster with titanium grid was conducted. Some titanium lines were observed in the discharge chamber; however, the signals were too weak to estimate the erosion of the screen grid. Nevertheless, this technique appears to be the only non-intrusive real-time means to evaluate screen grid erosion, and improvement of the collection optics is proposed. Direct examination of the erosion species using laser-induced fluorescence (LIF) was determined to be the best method for a real-time accelerator grid erosion diagnostic. An approach for a quantitative LIF diagnostic was presented.

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

    SciTech Connect

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

    2006-01-15

    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{sup -4} Pa Xe (3.3x10{sup -6} Torr Xe) to 1.1x10{sup -3} Pa Xe (8.4x10{sup -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.

  4. NASA's Evolutionary Xenon Thruster (NEXT) Prototype Model 1R (PM1R) Ion Thruster and Propellant Management System Wear Test Results

    NASA Technical Reports Server (NTRS)

    VanNoord, Jonathan L.; Soulas, George C.; Sovey, James S.

    2010-01-01

    The results of the NEXT wear test are presented. This test was conducted with a 36-cm ion engine (designated PM1R) and an engineering model propellant management system. The thruster operated with beam extraction for a total of 1680 hr and processed 30.5 kg of xenon during the wear test, which included performance testing and some operation with an engineering model power processing unit. A total of 1312 hr was accumulated at full power, 277 hr at low power, and the remainder was at intermediate throttle levels. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The propellant management system performed without incident during the wear test. The ion engine and propellant management system were also inspected following the test with no indication of anomalous hardware degradation from operation.

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

    NASA Technical Reports Server (NTRS)

    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.

  6. Plasma Emission Characteristics From a High Current Hollow Cathode in an Ion Thruster Discharge Chamber

    NASA Astrophysics Data System (ADS)

    Foster, John E.; Patterson, Michael J.

    2002-11-01

    The presence of energetic ions produced by a hollow cathodes operating at high emission currents (greater than 5A) has been documented in the literature. In order to further elucidate these findings, an investigation of a high current cathode operating in an ion thruster discharge chamber has been undertaken. Using Langmuir probes, a low energy charged particle analyzer and emission spectroscopy, the behavior of the near-cathode plasma and the emitted ion energy distribution was characterized. The presence of energetic ions was confirmed. It was observed that these ions had energies in excess of the discharge voltage and thus cannot be simply explained by ions falling out of plasma through a potential difference of this order. Additionally, evidence provided by Langmuir probes suggests the existence of a double layer essentially separating the hollow cathode plasma column from the main discharge. The radial potential difference associated with this double layer was measured to be of order the ionization potential.

  7. Plasma Emission Characteristics from a High Current Hollow Cathode in an Ion Thruster Discharge Chamber

    NASA Technical Reports Server (NTRS)

    Foster, John E.; Patterson, Michael J.

    2002-01-01

    The presence of energetic ions produced by a hollow cathodes operating at high emission currents (greater than 5A) has been documented in the literature. In order to further elucidate these findings, an investigation of a high current cathode operating in an ion thruster discharge chamber has been undertaken. Using Langmuir probes, a low energy charged particle analyzer and emission spectroscopy, the behavior of the near-cathode plasma and the emitted ion energy distribution was characterized. The presence of energetic ions was confirmed. It was observed that these ions had energies in excess of the discharge voltage and thus cannot be simply explained by ions falling out of plasma through a potential difference of this order. Additionally, evidence provided by Langmuir probes suggests the existence of a double layer essentially separating the hollow cathode plasma column from the main discharge. The radial potential difference associated with this double layer was measured to be of order the ionization potential.

  8. Performance Characterization of Ion Thruster with Isolated Magnet Rings

    NASA Astrophysics Data System (ADS)

    Arthur, Neil; Foster, John

    2016-10-01

    Many ion sources use magnetic multipole confinement to increase the primary electron containment length. The magnetic circuit increases ion source efficiency and plasma density. A multipole source consists of rings or rows of magnets of alternating polarity. In multipole sources, the bulk of the discharge current is collected at the magnet surface, through the relatively narrow leak width. Ion engines for space propulsion are one application of multipole ion sources. Here we characterize a four ring, broad beam ion source under simulated beam extraction using typical performance metrics for ion engines while biasing the magnetic rings individually. By biasing the magnetic cusps, through isolated, conformal electrodes placed on the magnet rings, the current distribution to each individual cusp can be modified. The effect of ring bias on ion beam current, propellant utilization efficiency, and discharge losses is measured over a broad range of ring bias. Previous experiments have shown that the current distribution to the rings can be controlled, and this current distribution has tangible effects on the plasma properties and ion source operation. The goal is to gain insight into which magnetic ring current distributions will yield enhancements in engine performance.

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

    NASA Astrophysics Data System (ADS)

    Clemens, Adam; Burgess, David

    2016-03-01

    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.

  10. A hollow cathode neutralizer for a 30-cm diameter bombardment thruster

    NASA Technical Reports Server (NTRS)

    Bechtel, R. T.

    1973-01-01

    Recent improvements in overall thrustor performance have imposed new constraints on neutralizer performance. The use of compensated grid extraction system requires a re-evaluation of neutralizer position. In addition a suitable control logic for the neutralizer has proven difficult. A series of tests were conducted to determine what effect neutralizer cathode geometry has on performance. The parameters investigated included orifice diameter and length, and cathode diameter. Similar tests investigated open and enclosed keeper geometries. Neutralizer position tests with compensated grids suggest positions approximately 10 cm from the accelerator and radially out of the beam envelope should result in satisfactory performance and long life. Finally operation at keeper currents of 1.5 amp has resulted in lower total neutralizer power, the elimination of tip heater power, and suitable closed loop control of the neutralizer vaporizer.

  11. Power processor for a 20CM ion thruster

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Schoenfeld, A. D.; Cohen, E.

    1973-01-01

    A power processor breadboard for the JPL 20CM Ion Engine was designed, fabricated, and tested to determine compliance with the electrical specification. The power processor breadboard used the silicon-controlled rectifier (SCR) series resonant inverter as the basic power stage to process all the power to the ion engine. The breadboard power processor was integrated with the JPL 20CM ion engine and complete testing was performed. The integration tests were performed without any silicon-controlled rectifier failure. This demonstrated the ruggedness of the series resonant inverter in protecting the switching elements during arcing in the ion engine. A method of fault clearing the ion engine and returning back to normal operation without elaborate sequencing and timing control logic was evolved. In this method, the main vaporizer was turned off and the discharge current limit was reduced when an overload existed on the screen/accelerator supply. After the high voltage returned to normal, both the main vaporizer and the discharge were returned to normal.

  12. Status of the NEXT Ion Thruster Long-Duration Test After 10,100 hr and 207 kg Demonstrated

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

    2008-01-01

    The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to validate and qualify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the mission-derived throughput requirement of 300 kg. This wear test is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 21, 2007, the thruster has accumulated 10,100 hr of operation at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. The thruster has processed 207 kg of xenon and demonstrated a total impulse of 8.5 106 N-s; the highest total impulse ever demonstrated by an ion thruster in the history of space propulsion. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Overall ion thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. Lifetime-limiting component erosion rates have been consistent with the NEXT service life assessment, which predicts the earliest failure sometime after 750 kg of xenon propellant throughput; well beyond the mission-derived lifetime requirement. The NEXT wear test data confirm that the erosion of the discharge keeper orifice, enlarging of nominal-current-density accelerator grid aperture cusps, and the decrease in cold grid-gap observed during the NSTAR Extended Life Test have been mitigated. This paper presents the status of the NEXT LDT to date.

  13. Experimental validation of the dual positive and negative ion beam acceleration in the plasma propulsion with electronegative gases thruster

    SciTech Connect

    Rafalskyi, Dmytro Popelier, Lara; Aanesland, Ane

    2014-02-07

    The PEGASES (Plasma Propulsion with Electronegative Gases) thruster is a gridded ion thruster, where both positive and negative ions are accelerated to generate thrust. In this way, additional downstream neutralization by electrons is redundant. To achieve this, the thruster accelerates alternately positive and negative ions from an ion-ion plasma where the electron density is three orders of magnitude lower than the ion densities. This paper presents a first experimental study of the alternate acceleration in PEGASES, where SF{sub 6} is used as the working gas. Various electrostatic probes are used to investigate the source plasma potential and the energy, composition, and current of the extracted beams. We show here that the plasma potential control in such system is key parameter defining success of ion extraction and is sensitive to both parasitic electron current paths in the source region and deposition of sulphur containing dielectric films on the grids. In addition, large oscillations in the ion-ion plasma potential are found in the negative ion extraction phase. The oscillation occurs when the primary plasma approaches the grounded parts of the main core via sub-millimetres technological inputs. By controlling and suppressing the various undesired effects, we achieve perfect ion-ion plasma potential control with stable oscillation-free operation in the range of the available acceleration voltages (±350 V). The measured positive and negative ion currents in the beam are about 10 mA for each component at RF power of 100 W and non-optimized extraction system. Two different energy analyzers with and without magnetic electron suppression system are used to measure and compare the negative and positive ion and electron fluxes formed by the thruster. It is found that at alternate ion-ion extraction the positive and negative ion energy peaks are similar in areas and symmetrical in position with +/− ion energy corresponding to the amplitude of the applied

  14. Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion

    SciTech Connect

    Dey, Indranuj; Toyoda, Yuji; Yamamoto, Naoji; Nakashima, Hideki

    2015-12-15

    A miniature microwave electron cyclotron resonance plasma source [(discharge diameter)/(microwave cutoff diameter) < 0.3] has been developed at Kyushu University to be used as an ion thruster in micro-propulsion applications in the exosphere. The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency. Performance of the ion thruster is studied at 3 microwave frequencies (1.2 GHz, 1.6 GHz, and 2.45 GHz), for low input powers (<15 W) and small xenon mass flow rates (<40 μg/s), by experimentally measuring the extracted ion beam current through a potential difference of ≅1200 V. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.6 GHz. At this frequency, for an input power of 8 W, and propellant (xenon) mass flow rate of 21 μg/s, 13.7 mA of ion beam current is obtained, equivalent to an calculated thrust of 0.74 mN.

  15. Development of a miniature microwave electron cyclotron resonance plasma ion thruster for exospheric micro-propulsion.

    PubMed

    Dey, Indranuj; Toyoda, Yuji; Yamamoto, Naoji; Nakashima, Hideki

    2015-12-01

    A miniature microwave electron cyclotron resonance plasma source [(discharge diameter)/(microwave cutoff diameter) < 0.3] has been developed at Kyushu University to be used as an ion thruster in micro-propulsion applications in the exosphere. The discharge source uses both radial and axial magnetostatic field confinement to facilitate electron cyclotron resonance and increase the electron dwell time in the volume, thereby enhancing plasma production efficiency. Performance of the ion thruster is studied at 3 microwave frequencies (1.2 GHz, 1.6 GHz, and 2.45 GHz), for low input powers (<15 W) and small xenon mass flow rates (<40 μg/s), by experimentally measuring the extracted ion beam current through a potential difference of ≅1200 V. The discharge geometry is found to operate most efficiently at an input microwave frequency of 1.6 GHz. At this frequency, for an input power of 8 W, and propellant (xenon) mass flow rate of 21 μg/s, 13.7 mA of ion beam current is obtained, equivalent to an calculated thrust of 0.74 mN.

  16. Integral electrical characteristics and local plasma parameters of a RF ion thruster

    SciTech Connect

    Masherov, P. E.; Riaby, V. A.; Godyak, V. A.

    2016-02-15

    Comprehensive diagnostics has been carried out for a RF ion thruster based on inductively coupled plasma (ICP) source with an external flat antenna coil enhanced by ferrite core. The ICP was confined within a cylindrical chamber with low aspect ratio to minimize plasma loss to the chamber wall. Integral diagnostics of the ICP electrical parameters (RF power balance and coil current) allowed for evaluation of the antenna coils, matching networks, and eddy current loss and the true RF power deposited to plasma. Spatially resolved electron energy distribution functions, plasma density, electron temperatures, and plasma potentials were measured with movable Langmuir probes.

  17. Electric propulsion. [pulsed plasma thruster and electron bombardment ion engine for MSAT attitude control and stationkeeping

    NASA Technical Reports Server (NTRS)

    1982-01-01

    An alternative propulsion subsystem for MSAT is presented which has a potential of reducing the satellite weight by more than 15%. The characteristics of pulsed plasma and ion engines are described and used to estimate of the mass of the propellant and thrusters for attitude control and stationkeeping functions for MSAT. Preliminary estimates indicate that the electric propulsion systems could also replace the large momentum wheels necessary to counteract the solar pressure; however, the fine pointing wheels would be retained. Estimates also show that either electric propulsion system can save approximately 18% to 20% of the initial 4,000 kg mass. The issues that require further experimentation are mentioned.

  18. Integral electrical characteristics and local plasma parameters of a RF ion thruster.

    PubMed

    Masherov, P E; Riaby, V A; Godyak, V A

    2016-02-01

    Comprehensive diagnostics has been carried out for a RF ion thruster based on inductively coupled plasma (ICP) source with an external flat antenna coil enhanced by ferrite core. The ICP was confined within a cylindrical chamber with low aspect ratio to minimize plasma loss to the chamber wall. Integral diagnostics of the ICP electrical parameters (RF power balance and coil current) allowed for evaluation of the antenna coils, matching networks, and eddy current loss and the true RF power deposited to plasma. Spatially resolved electron energy distribution functions, plasma density, electron temperatures, and plasma potentials were measured with movable Langmuir probes.

  19. Impingement-Current-Erosion Characteristics of Accelerator Grids on Two-Grid Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Barker, Timothy

    1996-01-01

    Accelerator grid sputter erosion resulting from charge-exchange-ion impingement is considered to be a primary cause of failure for electrostatic ion thrusters. An experimental method was developed and implemented to measure erosion characteristics of ion-thruster accel-grids for two-grid systems as a function of beam current, accel-grid potential, and facility background pressure. Intricate accelerator grid erosion patterns, that are typically produced in a short time (a few hours), are shown. Accelerator grid volumetric and depth-erosion rates are calculated from these erosion patterns and reported for each of the parameters investigated. A simple theoretical volumetric erosion model yields results that are compared to experimental findings. Results from the model and experiments agree to within 10%, thereby verifying the testing technique. In general, the local distribution of erosion is concentrated in pits between three adjacent holes and trenches that join pits. The shapes of the pits and trenches are shown to be dependent upon operating conditions. Increases in beam current and the accel-grid voltage magnitude lead to deeper pits and trenches. Competing effects cause complex changes in depth-erosion rates as background pressure is increased. Shape factors that describe pits and trenches (i.e. ratio of the average erosion width to the maximum possible width) are also affected in relatively complex ways by changes in beam current, ac tel-grid voltage magnitude, and background pressure. In all cases, however, gross volumetric erosion rates agree with theoretical predictions.

  20. Particle Simulation of a Micro ICP Plasma Source for Miniature Ion Thruster

    NASA Astrophysics Data System (ADS)

    Takao, Yoshinori; Eriguchi, Koji; Ono, Kouich

    2009-10-01

    There has recently been an ongoing trend toward decreasing the mass, dimension, and overall complexity of spacecraft. Propulsion systems are no exception. We have developed an electrothermal-type microthruster so far, which can produce a relatively high thrust, and have investigated the thrust performance with an experimental and numerical approach. On the other hand, a microthruster with a high specific impulse, such as ion thruster, is also required. The micro ion thruster presented here uses a cylindrical micro ICP with a flat spiral coil for its ion source, the inner radius and the length of which are 3 mm and 6 mm, respectively. To investigate the plasma characteristics of the source, we have developed a particle simulation model (PIC/MC: Particle-in-Cell/Monte Carlo) for Ar gas as a propellant. The simulation results showed that the electron density obtained was ˜10^17 m^ -3 at an Ar gas pressure of 4 mTorr with an absorbed power of 10 mW, producing a thrust of 50 μN and specific impulse of 7000 s.

  1. High Power Ion Heating in Helium and Hydrogen Plasmas for Advanced Plasma Thrusters

    NASA Astrophysics Data System (ADS)

    Ando, Akira; Hagiwara, Tatsuya; Domon, Masakazu; Taguchi, Takahiro

    High power ion cyclotron resonance heating is performed in a fast-flowing plasma operated with hydrogen and helium gases. Ion heating is clearly observed in hydrogen plasma as well as in helium plasma. The resonance region of magnetic field is broader and wave absorption efficiency is higher in hydrogen plasma than those in helium plasma. The thermal energy of the heated ions is converted to the kinetic energy of the exhaust plume by passing through a diverging magnetic nozzle set in a downstream region. In the magnetic nozzle energy conversion occurred as keeping the magnetic moment constant, but some discrepancy was observed in larger gradient of magnetic field. The kinetic energy of the exhaust plume is successfully controlled by an input power of radio-frequency wave, which is one of the key technologies for the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) type plasma thruster.

  2. Measurement of ion thruster exhaust characteristics and interaction with simulated ATS-F spacecraft

    NASA Technical Reports Server (NTRS)

    Worlock, R.; Trump, G.; Sellen, J. M., Jr.; Kemp, R. F.

    1973-01-01

    The ATS-F ion engine was mounted on a simulated spacecraft and was operated in a 22 by 35 foot vacuum chamber, using the same neutralizer control point as in earlier small chamber tests. The control point was in the middle of a range of 16 steps and, thus, the range should be adequate for transition to space flight. Measurement of the near- and far-field ions showed that the ion beam was well defined in a cone of 18-degrees half-angle. The material deposition experiment indicated that the ATS-F solar array would accumulate less than 0.2 A of aluminum per thousand hours of thruster operation, so that the corresponding power loss could be considered negligible. An interesting result was that the coupling between the beam and spacecraft was strong enough to require relatively large increases in the beam potential as the neutralizer bias was increased.

  3. A retarding potential analyzer design for keV-level ion thruster beams

    NASA Astrophysics Data System (ADS)

    Zhang, Zhe; Tang, Haibin; Zhang, Zun; Wang, Joseph; Cao, Shuai

    2016-12-01

    We present a new Retarding Potential Analyzer (RPA) design that is capable of measuring keV-level energy, high-density plasma beams. This instrument overcomes the limitations of existing RPAs and can operate in plasmas with densities in excess of 1 × 1015 m-3 and ion energies up to 1200 eV. The RPA design parameters were determined by analyzing the electron density and temperature, the sheath thickness, and the ion density in the beam based on the Faraday probe and Langmuir probe measurements. A previously unobserved grid spacing arcing phenomenon was observed in experiments. This arcing phenomenon was also investigated and a grid spacing criterion was proposed to eliminate the arcing. We present measurement results on the plasma beam emitted from the 20 cm Xenon ion thruster used on the Chinese SJ-9A satellite.

  4. Observation of Ion Cyclotron Heating in a Fast-flowing Plasma for an Advanced Plasma Thruster

    NASA Astrophysics Data System (ADS)

    Ando, Akira; Hatanaka, Motoi; Shibata, Masaki; Tobari, Hiroyuki; Hattori, Kunihiko; Inutake, Masaaki

    2004-11-01

    In the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) project in NASA, the combined system of the ion cyclotron heating and the magnetic nozzle is proposed to control a ratio of specific impulse to thrust at constant power. In order to establish the advanced plasma thruster, experiments of an ion heating and plasma acceleration by a magnetic nozzle are performed in a fast-flowing plasma in the HITOP device. A fast-flowing He plasma is produced by Magneto-Plasma-Dynamic Arcjet (MPDA) operated with an externally-applied magnetic field up to 1kG. RF waves with an ion cyclotron range of frequency (f=20-300kHz) is excited by a helically-wound antenna located downstream of the MPDA. Increases of an ion temperature and plasma stored energy measured by a diamagnetic coil clearly observed during the RF pulse. The heating efficiency is compared for various magnetic field configurations and strengths. There appears no indication of cyclotron resonance in a high density plasma where the ratio of ion cyclotron frequency to ion-ion collision one is below unity, because an ion-ion collisional effect is dominant. When the density becomes low and the ratio of ion cyclotron frequency to ion-ion collision one becomes high, features of ion cyclotron resonance are clearly appeared. The optimum magnetic field strength for the ion heating is slightly lower than that of the cyclotron resonance, which is caused by the Doppler effect due to the fast-flowing plasma. An ion energy distribution function is measured at a magnetic nozzle region by an electrostatic analyzer and increase of the parallel velocity is also observed.

  5. A HiPIMS plasma source with a magnetic nozzle that accelerates ions: application in a thruster

    NASA Astrophysics Data System (ADS)

    Bathgate, Stephen N.; Ganesan, Rajesh; Bilek, Marcela M. M.; McKenzie, David R.

    2017-01-01

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

  6. Phase-resolved emission spectroscopy of a neutraliser-free gridded ion thruster

    NASA Astrophysics Data System (ADS)

    Dedrick, James; Gibson, Andrew; Rafalskyi, Dmytro; Aanesland, Ane

    2015-09-01

    Power-efficient electric propulsion systems that operate without an external neutraliser have the potential to increase the longevity of traditional concepts. The Neptune gridded-ion thruster prototype, which uses a single radio-requency (rf) power source for plasma generation, ion acceleration and beam neutralisation, is under development. Previous research has suggested that the time-resolved electron dynamics in the plume are important for maintaining charge neutrality and overall performance. In this study, the electron dynamics in the exhaust beam are investigated within the rf cycle using phase-resolved emission spectroscopy. The results are compared with time-resolved and time-integrated electrical diagnostics to investigate the mechanisms behind beam neutralisation. This work received financial support from the York-Paris CIRC and state aid managed by the laboratory of excellence Plas@Par (ANR-11-IDEX-0004-02).

  7. Ion Velocity and Plasma Potential Measurements of a Cylindrical Cusped Field Thruster

    DTIC Science & Technology

    2011-05-19

    Thrusters. Journal of Propulsion and Power. 1998;14(5):736–743. [4] Raitses Y, Fisch NJ. Parametric investigations of a nonconventional Hall thruster...Astronautics; 2009. . [7] Raitses Y, Granstedt E, Smirnov A, Merino E, Fisch NJ. Effects of Cathode Electron Emission on Hall Thruster Discharge. In...Proceedings of the 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit; 2008. . [8] Raitses Y, Smirnov A, Fisch NJ. Cylindrical Hall Thrusters. In

  8. Note: An advanced in situ diagnostic system for characterization of electric propulsion thrusters and ion beam sources

    NASA Astrophysics Data System (ADS)

    Bundesmann, C.; Tartz, M.; Scholze, F.; Leiter, H. J.; Scortecci, F.; Gnizdor, R. Y.; Neumann, H.

    2010-04-01

    We present an advanced diagnostic system for in situ characterization of electric propulsion thrusters and ion beam sources. The system uses a high-precision five-axis positioning system with a modular setup and the following diagnostic tools: a telemicroscopy head for optical imaging, a triangular laser head for surface profile scanning, a pyrometer for temperature scanning, a Faraday probe for current density mapping, and an energy-selective mass spectrometer for beam characterization (energy and mass distribution, composition). The capabilities of our diagnostic system are demonstrated with a Hall effect thruster SPT-100D EM1.

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

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Shastry, Rohit

    2016-01-01

    A Long Duration Test (LDT) was initiated in June 2005 as a part of NASA's 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. The post-test inspection objectives for the ion optics were derived from the original NEXT LDT test objectives, such as service life model validation, and expanded to encompass other goals that included verification of in situ measurements, test issue root causes, and past design changes. The ion optics cold grid gap had decreased only by an average of 7% of pretest center grid gap, so efforts to stabilize NEXT grid gap were largely successful. The upstream screen grid surface exhibited a chamfered erosion pattern. Screen grid thicknesses were = 86% of the estimated pretest thickness, indicating that the screen grid has substantial service life remaining. Deposition was found on the screen aperture walls and downstream surfaces that was primarily composed of grid material and back-sputtered carbon, and this deposition likely caused the minor decreases in screen grid ion transparency during the test. Groove depths had eroded through up to 35% of the accelerator grid thickness. Minimum accelerator aperture diameters increased only by about 5-7% of the pretest values and downstream surface diameters increased by about 24-33% of the pretest diameters. These results suggest that increasing the accelerator aperture diameters, improving manufacturing tolerances, and masking down the perforated diameter to 36 cm were successful in reducing the degree of accelerator aperture erosion at larger radii.

  10. Post-Test Analysis of the Deep Space One Spare Flight Thruster Ion Optics

    NASA Technical Reports Server (NTRS)

    Anderson, John R.; Sengupta, Anita; Brophy, John R.

    2004-01-01

    The Deep Space 1 (DSl) spare flight thruster (FT2) was operated for 30,352 hours during the extended life test (ELT). The test was performed to validate the service life of the thruster, study known and identify unknown life limiting modes. Several of the known life limiting modes involve the ion optics system. These include loss of structural integrity for either the screen grid or accelerator grid due to sputter erosion from energetic ions striking the grid, sputter erosion enlargement of the accelerator grid apertures to the point where the accelerator grid power supply can no longer prevent electron backstreaming, unclearable shorting between the grids causes by flakes of sputtered material, and rouge hole formation due to flakes of material defocusing the ion beam. Grid gap decrease, which increases the probability of electron backstreaming and of arcing between the grids, was identified as an additional life limiting mechanism after the test. A combination of accelerator grid aperture enlargement and grid gap decrease resulted in the inability to prevent electron backstreaming at full power at 26,000 hours of the ELT. Through pits had eroded through the accelerator grid webbing and grooves had penetrated through 45% of the grid thickness in the center of the grid. The upstream surface of the screen grid eroded in a chamfered pattern around the holes in the central portion of the grid. Sputter deposited material, from the accelerator grid, adhered to the downstream surface of the screen grid and did not spall to form flakes. Although a small amount of sputter deposited material protruded into the screen grid apertures, no rouge holes were found after the ELT.

  11. Implementation and verification of a hybrid performance and impedance model of gridded radio-frequency ion thrusters

    NASA Astrophysics Data System (ADS)

    Volkmar, Chris; Ricklefs, Ubbo

    2015-10-01

    In this paper, we show the development steps for an iterative performance and impedance model of gridded radio-frequency (RF) ion thrusters. The input parameters are equivalent to those of the real propulsion system; i.e., coil current, propellant mass flow, and extraction grid voltages. Therefore, the model is easily validated and verified by experimental data and can furthermore be used to optimize the overall thruster performance. The model predicts volume-averaged plasma parameters such as electron temperature, conductivity, total pressure, and ionization fraction as well as thruster performance data like generated thrust, specific impulse, and mass and electrical efficiency. The above mentioned plasma parameters are obtained as functions of the discharge chamber's geometry by using a charge balance equation that relates generated ions to ions lost at the chamber's walls. The plasma related quantities influence the electromagnetic field penetration which is here evaluated by means of a diffusion equation for the vector potential. The vector potential is obtained by a 3D Finite-Difference-Method on a cubic and rectangular grid which, in principle, offers the opportunity to have arbitrary plasma chamber and coil geometries. An actual ion thruster geometry is evaluated in this study in favor of experimental verification of the numerically obtained data. The thruster's coil generates highly asymmetric electromagnetic fields which motivates the use of a three-dimensional solver. A dissipation model based on Ohm's law and Poynting's theorem is used to determine the absorbed power within the discharge. To obtain a stable solution, the electromagnetically absorbed power is equated to the power lost due to elastic and inelastic collisions and electron wall flux. This whole process is iteratively repeated until the degree of ionization converges within a given threshold. To relate the stable discharge parameters to the thruster performance an extraction model based on a

  12. Closed Loop solar array-ion thruster system with power control circuitry

    NASA Technical Reports Server (NTRS)

    Gruber, R. P. (Inventor)

    1979-01-01

    A power control circuit connected between a solar array and an ion thruster receives voltage and current signals from the solar array. The control circuit multiplies the voltage and current signals together to produce a power signal which is differentiated with respect to time. The differentiator output is detected by a zero crossing detector and, after suitable shaping, the detector output is phase compared with a clock in a phase demodulator. An integrator receives no output from the phase demodulator when the operating point is at the maximum power but is driven toward the maximum power point for non-optimum operation. A ramp generator provides minor variations in the beam current reference signal produced by the integrator in order to obtain the first derivative of power.

  13. Optimized electrode placement along the channel of a Hall thruster for ion focusing

    SciTech Connect

    Qing, Shaowei; E, Peng; Xia, Guangqing; Tang, Ming-Chun; Duan, Ping

    2014-01-21

    An optimal placement of the segmented electrode for increasing the lifetime of the Aton-type Hall thruster, i.e., reducing the plume divergence, is demonstrated using a 2D3V fully kinetic Particle-in-Cell method. Segmented electrodes, embedded near the ionization region of non-segmented case and biased above anode potential, lead to an increased separation between the ionization and acceleration regions and the formation of an efficient acceleration electric field configuration as potential lens. Due to this electrode placement, the sheath near the ceramic walls of the acceleration region is collapsed and an excellent ion beam focusing is demonstrated. The potential contour pockets around the electrodes and the sheath collapse phenomenon are also discussed.

  14. Solutions for discharge chamber sputtering and anode deposit spalling in small mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Power, J. L.; Hiznay, D. J.

    1975-01-01

    Proposed solutions to the problems of sputter erosion and sputtered material spalling in the discharge chamber of small mercury ion thrusters are presented. The accelerated life test evaluated three such proposed solutions: (1) the use of tantalum as a single low sputter yield material for the exposed surfaces of the discharge chamber components subject to sputtering, (2) the use of a severely roughened anode surface to improve the adhesion of the sputter-deposited coating, and (3) the use of a wire cloth anode surface in order to limit the size of any coating flakes which might spall from it. Because of the promising results obtained in the accelerated life test with anode surfaces roughened by grit-blasting, experiments were carried out to optimize the grit-blasting procedure. The experimental results and an optimal grit-blasting procedure are presented.

  15. Pulse ignition characterization of mercury ion thruster hollow cathode using an improved pulse ignitor

    NASA Technical Reports Server (NTRS)

    Wintucky, E. G.; Gruber, R. P.

    1978-01-01

    An investigation of the high voltage pulse ignition characteristics of the 8 cm mercury ion thruster neutralizer cathode identified a low rate of voltage rise and long pulse duration as desirable factors for reliable cathode starting. Cathode starting breakdown voltages were measured over a range of mercury flow rates and tip heater powers for pulses with five different rates of voltage rise. Breakdown voltage requirements for the fastest rising pulse (2.5 to 3.0 kV/micro sec) were substantially higher (2 kV or more) than for the slowest rising pulse (0.3 to 0.5 kV/micro sec) for the same starting conditions. Also described is an improved, low impedance pulse ignitor circuit which reduces power losses and eliminates problems with control and packaging associated with earlier designs.

  16. Argon hollow cathode. M.S. Thesis; [propellants for ion bombardment thrusters

    NASA Technical Reports Server (NTRS)

    Rehn, L. A.

    1976-01-01

    An interest in alternate propellants for ion-bombardment thrusters, together with ground applications of this technology, has prompted consideration of argon. Several variations of conventional hollow cathode designs were tried, but the bulk of the testing used a hollow tube with an internal tungsten emitter and an orifice at one end. The optimum cathode tube diameter was found to be in the range of 1.0-2.5 cm, somewhat larger than those used for cesium and mercury. Optimum orifice diameter depended on operating conditions, and varied from 0.5 to 5 mm. Biasing the internal emitter negative relative to the cathode chamber reduced the external coupling voltage and should therefore improve orifice lifetime. The expected effect of this bias on emitter lifetime was less clear. Lifetime tests were not conducted as part of this investigation, but several designs show promise of long lifetime in specific applications.

  17. Computed versus measured ion velocity distribution functions in a Hall effect thruster

    SciTech Connect

    Garrigues, L.; Mazouffre, S.; Bourgeois, G.

    2012-06-01

    We compare time-averaged and time-varying measured and computed ion velocity distribution functions in a Hall effect thruster for typical operating conditions. The ion properties are measured by means of laser induced fluorescence spectroscopy. Simulations of the plasma properties are performed with a two-dimensional hybrid model. In the electron fluid description of the hybrid model, the anomalous transport responsible for the electron diffusion across the magnetic field barrier is deduced from the experimental profile of the time-averaged electric field. The use of a steady state anomalous mobility profile allows the hybrid model to capture some properties like the time-averaged ion mean velocity. Yet, the model fails at reproducing the time evolution of the ion velocity. This fact reveals a complex underlying physics that necessitates to account for the electron dynamics over a short time-scale. This study also shows the necessity for electron temperature measurements. Moreover, the strength of the self-magnetic field due to the rotating Hall current is found negligible.

  18. Transport of ion beam in an annular magnetically expanding helicon double layer thruster

    SciTech Connect

    Zhang, Yunchao Charles, Christine; Boswell, Rod

    2014-06-15

    An ion beam generated by an annular double layer has been measured in a helicon thruster, which sustains a magnetised low-pressure (5.0 × 10{sup −4} Torr) argon plasma at a constant radio-frequency (13.56 MHz) power of 300 W. After the ion beam exits the annular structure, it merges into a solid centrally peaked structure in the diffusion chamber. As the annular ion beam moves towards the inner region in the diffusion chamber, a reversed-cone plasma wake (with a half opening angle of about 30°) is formed. This process is verified by measuring both the radial and axial distributions of the beam potential and beam current. The beam potential changes from a two-peak radial profile (maximum value ∼ 30 V, minimum value ∼ 22.5 V) to a flat (∼28 V) along the axial direction; similarly, the beam current changes from a two-peak to one-peak radial profile and the maximum value decreases by half. The inward cross-magnetic-field motion of the beam ions is caused by a divergent electric field in the source. Cross-field diffusion of electrons is also observed in the inner plume and is determined as being of non-ambipolar origin.

  19. The Investigation Of Carbon Contamination And Sputtering Effects Of Xenon Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Prak, Moline K.

    2004-01-01

    The Electro-Physics Branch of the NASA Glenn Research Center investigates the effect of atomic oxygen, environmental durability of high performance power materials and surfaces, and low earth orbit. One of its current projects involves the analysis of ion thrusters. Ion thrusters are devices that initiate a beam of ions to a target area. The type of ion thruster that I have been working with this Summer of 2004 emits positively charged Xenon (Xe(+)) atoms through two grids, the screen grid and the accelerator grid, after it enters an ionization chamber. Insulators are used to mechanically hold and separate these two grids. A propellant isolator, an instrument that closely resembles insulators, is placed in front of the ionization chamber. Both the insulator and isolator are made with a ceramic compound and filled with insulating beads. The main difference between the two devices is that the propellant isolator allows gas to flow through, in this case, the gas is Xe(+) and the insulators do not. In order to avoid carbon deposits and other contaminating chemicals to settle on the insulators and propellant isolator, a metal shadow shield is placed around them. These shadow shields function as a protectant and can be shaped in numerous configurations. Part of my job responsibility this summer is to investigate the effectiveness of different shadow shields that are utilized on three different ion engines: the NSTAR (NASA Solar Electric Propulsion Technology Application Readiness), JIMO (Jupiter Icy Moons Orbiter), and NEXIS (Nuclear Electric Xenon Ion System). Using calculus and other mathematical tactics, I was asked to find the total flux of carbon contamination that was able to pass the protectant shadow shield. I familiarized myself with the software program, MathCad2004, to help perform some mathematical computations such as complex integration. Another method of studying the probability of contamination is by experimental simulation. After attaining the precise

  20. Feasibility Study on Performance Enhancement Options for the ECR Ion Thruster µ10

    NASA Astrophysics Data System (ADS)

    Nishiyama, Kazutaka; Hosoda, Satoshi; Usui, Miyuki; Tsukizaki, Ryudo; Hayashi, Hiroshi; Shimizu, Yukio; Kuninaka, Hitoshi

    In order to adapt to a wide variety of the space flights, such as small geosynchronous satellites and deep space explorers, feasibility study of performance enhancement options for the microwave discharge ion thruster µ10 is underway. Authors are considering the following five options: 1. Lower insertion loss DC blocks; 2. Direct monopole antenna insertion to the discharge chamber without using a circular waveguide part; 3. Optimization of gas injector layout which was originally located deep in the waveguide; 4. Additional magnet rings aiming ion loss reduction to the side wall of the discharge chamber; 5. New ion optics consists of a thinner screen grid and a smaller-hole accelerator grid. Not all but most of them have already been tested and reported in this article. The original models for Hayabusa asteroid explorer generated 8 mN at maximum. Larger thrust generation was impossible even if propellant flow rates and microwave powers were increased. It turned out to be feasible to increase the maximum thrust to a range of 10 - 11 mN with above mentioned options by supplying more flow rates and/or more microwave powers.

  1. Interaction of a solar array with an ion thruster due to the charge-exchange plasma

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1976-01-01

    The generation of a charge exchange plasma by a thruster, the transport of this plasma to the solar array, and the interaction of the solar array with the plasma after it arrives are all described. The generation of this plasma is described accurately from thruster geometry and operating conditions. The transport of the charge exchange plasma was studied experimentally with a 15 cm thruster. A model was developed for simple thruster array configurations. A variety of experiments were surveyed for the interaction of the plasma at the solar array.

  2. Multipole gas thruster design. Ph.D. Thesis

    NASA Technical Reports Server (NTRS)

    Isaacson, G. C.

    1977-01-01

    The development of a low field strength multipole thruster operating on both argon and xenon is described. Experimental results were obtained with a 15-cm diameter multipole thruster and are presented for a wide range of discharge-chamber configurations. Minimum discharge losses were 300-350 eV/ion for argon and 200-250 eV/ion for xenon. Ion beam flatness parameters in the plane of the accelerator grid ranged from 0.85 to 0.93 for both propellants. Thruster performance is correlated for a range of ion chamber sizes and operating conditions as well as propellant type and accelerator system open area. A 30-cm diameter ion source designed and built using the procedure and theory presented here-in is shown capable of low discharge losses and flat ion-beam profiles without optimization. This indicates that by using the low field strength multipole design, as well as general performance correlation information provided herein, it should be possible to rapidly translate initial performance specifications into easily fabricated, high performance prototypes.

  3. A 15,000-hour cyclic endurance test of an 8-centimeter-diameter electron bombardment mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Nakanishi, S.

    1976-01-01

    A laboratory model 8 cm thruster with improvements to minimize ion chamber erosion and peeling of sputtered metal was subjected to a cyclic endurance test for 15,040 hours and 460 restarts. A charted history of several thruster operating variables and off-normal events are shown in 600-hour segments at three points in the test. The transient behavior of these variables during a typical start-stop cycle is presented. Finding of the post-test inspection confirmed most of the expected results. Charge exchange ions caused normal accelerator grid erosion. The workability of the various design features was substantiated, and attainable improvements in propellant utilization efficiency should significantly reduce accelerator erosion.

  4. Cylindrical geometry hall thruster

    DOEpatents

    Raitses, Yevgeny; Fisch, Nathaniel J.

    2002-01-01

    An apparatus and method for thrusting plasma, utilizing a Hall thruster with a cylindrical geometry, wherein ions are accelerated in substantially the axial direction. The apparatus is suitable for operation at low power. It employs small size thruster components, including a ceramic channel, with the center pole piece of the conventional annular design thruster eliminated or greatly reduced. Efficient operation is accomplished through magnetic fields with a substantial radial component. The propellant gas is ionized at an optimal location in the thruster. A further improvement is accomplished by segmented electrodes, which produce localized voltage drops within the thruster at optimally prescribed locations. The apparatus differs from a conventional Hall thruster, which has an annular geometry, not well suited to scaling to small size, because the small size for an annular design has a great deal of surface area relative to the volume.

  5. Optimization of Cylindrical Hall Thrusters

    SciTech Connect

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

    2007-07-24

    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. __________________________________________________

  6. Optimization of Cylindrical Hall Thrusters

    SciTech Connect

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

    2007-11-27

    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.

  7. Faraday Probe Analysis, Part 2: Evaluation of Facility Effects on Ion Migration in a Hall Thruster Plume (Preprint)

    DTIC Science & Technology

    2010-02-24

    and may alter the thruster performance and plume properties [1,2,3]. A number of investigations have studied nude Faraday probe design...Release, Distribution Unlimited (JA-2010-064) 3 propulsion plumes. One approach is to attach a collimator to the entrance of a nude Faraday probe to...energy CEX ions. A study by Rovey [5] compared results from a magnetically filtered Faraday probe, a boxed Faraday probe, and a nude Faraday probe to

  8. Development of Power Electronics for a 0.2kW-Class Ion Thruster

    NASA Technical Reports Server (NTRS)

    Pinero, Luis R.; Patterson, Michael J.; Bowers, Glen E.

    1997-01-01

    Applications that might benefit from low power ion propulsion systems include Earth-orbit magnetospheric mapping satellite constellations, low Earth-orbit satellites, geosynchronous Earth-orbit satellite north-south stationkeeping, and asteroid orbiters. These spacecraft are likely to have masses on the order of 50 to 500 kg with up to 0.5 kW of electrical power available. A power processing unit for a 0.2 kW-class ion thruster is currently under development for these applications. The first step in this effort is the development and testing of a 0.24 kW beam power supply. The design incorporates a 20 kHz full bridge topology with multiple secondaries connected in series to obtain outputs of up to 1200 V(sub DC). A current-mode control pulse width modulation circuit built using discrete components was selected for this application. An input voltage of 28 +/- 4 V(sub DC) was assumed, since the small spacecraft for which this system is targeted are anticipated to have unregulated low voltage busses. Efficiencies in excess of 91 percent were obtained at maximum output power. The total mass of the breadboard was less than 1.0 kg and the component mass was 0.53 kg. It is anticipated that a complete flight power processor could weigh about 2.0 kg.

  9. Measurement of xenon plasma properties in an ion thruster using laser Thomson scattering technique

    SciTech Connect

    Yamamoto, N.; Tomita, K.; Sugita, K.; Kurita, T.; Nakashima, H.; Uchino, K.

    2012-07-15

    This paper reports on the development of a method for measuring xenon plasma properties using the laser Thomson scattering technique, for application to ion engine system design. The thresholds of photo-ionization of xenon plasma were investigated and the number density of metastable atoms, which are photo-ionized by a probe laser, was measured using laser absorption spectroscopy, for several conditions. The measured threshold energy of the probe laser using a plano-convex lens with a focal length of 200 mm was 150 mJ for a xenon mass flow rate of 20 {mu}g/s and incident microwave power of 6 W; the probe laser energy was therefore set as 80 mJ. Electron number density was found to be (6.2 {+-} 0.4) Multiplication-Sign 10{sup 17} m{sup -3} and electron temperature was found to be 2.2 {+-} 0.4 eV at a xenon mass flow rate of 20 {mu}g/s and incident microwave power of 6 W. The threshold of the probe laser intensity against photo-ionization in a miniature xenon ion thruster is almost constant for various mass flow rates, since the ratio of population of the metastable atoms to the electron number density is little changed.

  10. Evaluation of High-Power Solar Electric Propulsion using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo Missions

    NASA Technical Reports Server (NTRS)

    Frisbee, Robert H.

    2006-01-01

    This paper presents the results of mission analyses that expose the advantages and disadvantages of high-power (MWe-class) Solar Electric Propulsion (SEP) for Lunar and Mars Cargo missions that would support human exploration of the Moon and Mars. In these analyses, we consider SEP systems using advanced Ion thrusters (the Xenon [Xe] propellant Herakles), Hall thrusters (the Bismuth [Bi] propellant Very High Isp Thruster with Anode Layer [VHITAL], magnetoplasmadynamic (MPD) thrusters (the Lithium [Li] propellant Advanced Lithium-Fed, Applied-field Lorentz Force Accelerator (ALFA2), and pulsed inductive thruster (PIT) (the Ammonia [NH3] propellant Nuclear-PIT [NuPIT]). The analyses include comparison of the advanced-technology propulsion systems (VHITAL, ALFA2, and NuPIT) relative to state-of-theart Ion (Herakles) propulsion systems and quantify the unique benefits of the various technology options such as high power-per-thruster (and/or high power-per-thruster packaging volume), high specific impulse (Isp), high-efficiency, and tankage mass (e.g., low tankage mass due to the high density of bismuth propellant). This work is based on similar analyses for Nuclear Electric Propulsion (NEP) systems.

  11. Electrical Prototype Power Processor for the 30-cm Mercury electric propulsion engine

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Frye, R. J.

    1978-01-01

    An Electrical Prototpye Power Processor has been designed to the latest electrical and performance requirements for a flight-type 30-cm ion engine and includes all the necessary power, command, telemetry and control interfaces for a typical electric propulsion subsystem. The power processor was configured into seven separate mechanical modules that would allow subassembly fabrication, test and integration into a complete power processor unit assembly. The conceptual mechanical packaging of the electrical prototype power processor unit demonstrated the relative location of power, high voltage and control electronic components to minimize electrical interactions and to provide adequate thermal control in a vacuum environment. Thermal control was accomplished with a heat pipe simulator attached to the base of the modules.

  12. Observation of a high-energy tail in ion energy distribution in the cylindrical Hall thruster plasma

    SciTech Connect

    Lim, Youbong; Kim, Holak; Choe, Wonho Lee, Seung Hun; Seon, Jongho; Lee, Hae June

    2014-10-15

    A novel method is presented to determine populations and ion energy distribution functions (IEDFs) of individual ion species having different charge states in an ion beam from the measured spectrum of an E × B probe. The inversion of the problem is performed by adopting the iterative Tikhonov regularization method with the characteristic matrices obtained from the calculated ion trajectories. In a cylindrical Hall thruster plasma, an excellent agreement is observed between the IEDFs by an E × B probe and those by a retarding potential analyzer. The existence of a high-energy tail in the IEDF is found to be mainly due to singly charged Xe ions, and is interpreted in terms of non-linear ion acceleration.

  13. Helicon Plasma Source and Ion Beam Creation Characteristics of the MadHex Thruster

    NASA Astrophysics Data System (ADS)

    Scharer, J.; Wiebold, M.; He, R.

    2009-12-01

    Non-invasive measurements are performed on a pulsed and steady-state argon helicon plasma thruster with a static axial magnetic nozzle field (1 kG source, 1.5 kG nozzle peak). The helicon wave propagation is closely related to whistler modes that propagate in the Earth's ionosphere. Flow rates obtained are from less than 1 to 30 sccm with coupled 13.56 MHz rf power levels of between 700 W and 10 kW. Ion beam acceleration from electric fields caused by neutral depletion and double layers (DLs) similar to those detected by satellites in the Earth's aurora are observed. Collisional-radiative (CR) models for Ar II and Ar I are used to spectroscopically determine the electron temperature (Te) and the neutral density, respectively. The electron density (nemax=8 x 10^13/cc) is measured via 105 GHz microwave interferometry (IF) and is an input to the CR models. In collisionless, highly neutral-depleted regions, Te rises linearly with power while ne remains constrained. Regions of pressure balance and pressure gradients are present, and evidence of substantial axially accelerated ion flows is observed. Regimes where cooler (5 eV) and hotter (>20 eV) electron temperatures are observed for lower and higher flow rates. The axial ion energy distribution function and its acceleration is measured from the helicon source region thru the magnetic nozzle using tunable diode laser-induced fluorescence (LIF). We will present results of RF creation and optimization of thermal- and hot-electron components to enhance the thrust of the helicon double layer and discuss the character of the ion beam distribution as it moves through the DL region. The experiment will optimize rf power, mass flow rate, magnetic field, and helicon dynamic frequency with LIF, mm wave IF diagnostic measurements. A description of the ion acceleration process that has potential applications for spacecraft propulsion and is related to ion acceleration processes observed in the Earth's aurora will be discussed.

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

    SciTech Connect

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

    2015-02-15

    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.

  15. Ion Velocity Distribution in a Low-Power Cylindrical Hall Thruster

    DTIC Science & Technology

    2010-07-01

    S. A. and Fisch , N. J., “Cylindrical Hall Thrusters,” Proceedings of the 37th AIAA Plasmadynamics and Lasers Conference, No. AIAA-2006-3245, American...Channel of Low-Power Hall Thruster,” IEEE Transactions on Plasma Science, Vol. 36, No. 5, October 2008, pp. 1989–1997. 4Raitses, Y., S. A. and Fisch , N. J... Fisch , N. J., “Enhanced Ionization in the Cylindrical Hall Thruster,” Journal of Applied Physics, Vol. 94, No. 2, 2003. 16Gildea, S. R., B. O. and

  16. Laser characterization of the unsteady 2-D ion flow field in a Hall thruster with breathing mode oscillations

    NASA Astrophysics Data System (ADS)

    Lucca Fabris, Andrea; Young, Christopher; MacDonald-Tenenbaum, Natalia; Hargus, William, Jr.; Cappelli, Mark

    2016-10-01

    Hall thrusters are a mature form of electric propulsion for spacecraft. One commonly observed low frequency (10-50 kHz) discharge current oscillation in these E × B devices is the breathing mode, linked to a propagating ionization front traversing the channel. The complex time histories of ion production and acceleration in the discharge channel and near-field plume lead to interesting dynamics and interactions in the central plasma jet and downstream plume regions. A time-resolved laser-induced fluorescence (LIF) diagnostic non-intrusively measures 2-D ion velocity and relative ion density throughout the plume of a commercial BHT-600 Hall thruster manufactured by Busek Co. Low velocity classes of ions observed in addition to the main accelerated population are linked to propellant ionization outside of the device. Effects of breathing mode dynamics are shown to persist far downstream where modulations in ion velocity and LIF intensity are correlated with discharge current oscillations. This work is sponsored by the U.S. Air Force Office of Scientific Research with Dr. M. Birkan as program manager. C.Y. acknowledges support from the DOE NSSA Stewardship Science Graduate Fellowship under contract DE-FC52-08NA28752.

  17. The 5200 cycle test of an 8-cm diameter Hg ion thruster

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.; Wintucky, E. G.

    1978-01-01

    An accelerated cycle test was conducted in which an 8-cm Engineering Model Thruster (EMT) prototype successfully completed 5200 on-off cycles and a total of more than 1300hours of thruster operation at a 4.5 mN thrust level. Cathode tip heater powers required for starting and keeper voltages remained well within acceptable limits. The discharge chamber utilization and electrical efficiency were nearly constant over the duration of the test. It was concluded that on-off cyclic operation by itself does not appreciably degrade starting capability or performance of the 8-cm EMT.

  18. Experimental Demonstration of Microwave Signal/Electric Thruster Plasma Interaction Effects

    NASA Technical Reports Server (NTRS)

    Zaman, Afroz J.; Lambert, Kevin M.; Curran, Frank M.

    1995-01-01

    An experiment was designed and conducted in the Electric Propulsion Laboratory of NASA Lewis Research Center to assess the impact of ion thruster exhaust plasma plume on electromagnetic signal propagation. A microwave transmission experiment was set up inside the propulsion test bed using a pair of broadband horn antennas and a 30 cm 2.3 kW ion thruster. Frequency of signal propagation covered from 6.5 to 18 GHz range. The stainless steel test bed when enclosed can be depressurized to simulate a near vacuum environment. A pulsed CW system with gating hardware was utilized to eliminate multiple chamber reflections from the test signal. Microwave signal was transmitted and received between the two hours when the thruster was operating at a given power level in such a way that the signal propagation path crossed directly through the plume volume. Signal attenuation and phase shift due to the plume was measured for the entire frequency band. Results for this worst case configuration simulation indicate that the effects of the ion thruster plume on microwave signals is a negligible attenuation (within 0.15 dB) and a small phase shift (within 8 deg.). This paper describes the detailed experiment and presents some of the results.

  19. Enhanced Performance of Cylindrical Hall Thrusters

    SciTech Connect

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

    2007-05-14

    The cylindrical thruster differs significantly in its underlying physical mechanisms from the conventional annular Hall thruster. It features high ionization efficiency, quiet operation, ion acceleration in a large volume-to-surface ratio channel, and performance comparable with the state-of-the-art conventional Hall thrusters. Very significant plume narrowing, accompanied by the increase of the energetic ion fraction and improvement of ion focusing, led to 50%–60% increase of the thruster anode efficiency. These improvements were achieved by overrunning the discharge current in the magnetized thruster plasma.

  20. 6 GHz Microwave Power-Beaming Demonstration with 6-kV Rectenna and Ion-Breeze Thruster

    NASA Astrophysics Data System (ADS)

    Cummings, T.; Janssen, J.; Karnesky, J.; Laks, D.; Santillo, M.; Strause, B.; Myrabo, L. N.; Alden, A.; Bouliane, P.; Zhang, M.

    2004-03-01

    On 14 April 2003 at the Communications Research Center (CRC) in Ottawa, Ontario, a 5.85-GHz transmitter beamed 3-kW of microwave power to a remote rectifying antenna (i.e., rectenna) that delivered 6-kV to a special `Ion-Breeze' Engine (IBE). Three of CRC's 26.5-cm by 31-cm rectennas were connected in series to provide the ~6-kV output. RPI's low-voltage IBE thrusters performed well in a ``world's first'' power-beaming demonstration with rectennas and endoatmospheric ion-propulsion engines. The successful tests were a low-tech, proof-of-concept demonstration for the future full-sized MicroWave Lightcraft (MWLC) and its air breathing `loiter' propulsion mode. Additional IBE experiments investigated the feasibility of producing flight control forces on the MWLC. The objective was to torque the charged hull for `pitch' or `roll' maneuvers. The torquing demonstration was entirely successful.

  1. Further study of the effect of the downstream plasma condition on accelerator grid erosion in an ion thruster

    NASA Technical Reports Server (NTRS)

    Peng, Xiaohang; Ruyten, Wilhelmus M.; Keefer, Dennis

    1992-01-01

    Further numerical results are presented of earlier particle-in-cell/Monte Carlo calculations of accelerator grid erosion in an ion thruster. A comparison between numerical and experimental results suggests that the accelerator grid impingement is primarily due to ions created far downstream from the accelerator grid. In particular, for the same experimental conditions as those of Monheiser and Wilbur at Colorado State University, it is found that a downstream plasma density of 2 x 10 exp 14/cu m is required to give the same ratio of accelerator grid impingement current to beam current (5 percent). For this condition, a potential hill is found in the downstream region of 2.5 V.

  2. Hall Thruster Technology for NASA Science Missions

    NASA Technical Reports Server (NTRS)

    Manzella, David; Oh, David; Aadland, Randall

    2005-01-01

    The performance of a prototype Hall thruster designed for Discovery-class NASA science mission applications was evaluated at input powers ranging from 0.2 to 2.9 kilowatts. These data were used to construct a throttle profile for a projected Hall thruster system based on this prototype thruster. The suitability of such a Hall thruster system to perform robotic exploration missions was evaluated through the analysis of a near Earth asteroid sample return mission. This analysis demonstrated that a propulsion system based on the prototype Hall thruster offers mission benefits compared to a propulsion system based on an existing ion thruster.

  3. Electric thruster research

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1982-01-01

    It has been customary to assume that ions flow nearly equally in all directions from the ion production region within an electron-bombardment discharge chamber. In general, the electron current through a magnetic field can alter the electron density, and hence the ion density, in such a way that ions tend to be directed away from the region bounded by the magnetic field. When this mechanism is understood, it becomes evident that many past discharge chamber designs have operated with a preferentially directed flow of ions. Thermal losses were calculated for an oxide-free hollow cathode. At low electron emissions, the total of the radiation and conduction losses agreed with the total discharge power. At higher emissions, though, the plasma collisions external to the cathode constituted an increasingly greater fraction of the discharge power. Experimental performance of a Hall-current thruster was adversely affected by nonuniformities in the magnetic field, produced by the cathode heating current. The technology of closed-drift thrusters was reviewed. The experimental electron diffusion in the acceleration channel was found to be within about a factor of 3 of the Bohm value for the better thruster designs at most operating conditions. Thruster efficiencies of about 0.5 appear practical for the 1000 to 2000 s range of specific impulse. Lifetime information is limited, but values of several thousands of hours should be possible with anode layer thrusters operated or = to 2000 s.

  4. Fast Camera Imaging of Hall Thruster Ignition

    SciTech Connect

    C.L. Ellison, Y. Raitses and N.J. Fisch

    2011-02-24

    Hall thrusters provide efficient space propulsion by electrostatic acceleration of ions. Rotating electron clouds in the thruster overcome the space charge limitations of other methods. Images of the thruster startup, taken with a fast camera, reveal a bright ionization period which settles into steady state operation over 50 μs. The cathode introduces azimuthal asymmetry, which persists for about 30 μs into the ignition. Plasma thrusters are used on satellites for repositioning, orbit correction and drag compensation. The advantage of plasma thrusters over conventional chemical thrusters is that the exhaust energies are not limited by chemical energy to about an electron volt. For xenon Hall thrusters, the ion exhaust velocity can be 15-20 km/s, compared to 5 km/s for a typical chemical thruster

  5. A mercury flow meter for ion thruster testing. [response time, thermal sensitivity

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1973-01-01

    The theory of operation of the thermal flow meter is presented, and a theoretical model is used to determine design parameters for a device capable of measuring mercury flows in the range of 0 to 5 gm/hr. Flow meter construction is described. Tests performed using a positive displacement mercury pump as well as those performed with the device in the feed line of an operating thruster are discussed. A flow meter response time of about a minute and a sensitivity of about 10 mv/gm/hr are demonstrated. Additional work to relieve a sensitivity of the device to variations in ambient temperature is indicated to improve its quantitative performance.

  6. Baffle aperture design study of hollow cathode equipped ion thrusters. M.S. Thesis Technical Report, 1 Dec. 1979 - 1 Oct. 1980

    NASA Technical Reports Server (NTRS)

    Brophy, J. R., Jr.; Wilbur, P. J.

    1980-01-01

    A simple theoretical model which can be used as an aid in the design of the baffle aperture region of a hollow cathode equipped ion thruster was developed. An analysis of the ion and electron currents in both the main and cathode discharge chambers is presented. From this analysis a model of current flow through the aperture, which is required as an input to the design model, was developed. This model was verified experimentally. The dominant force driving electrons through the aperture was the force due to the electrical potential gradient. The diffusion process was modeled according to the Bolm diffusion theory. A number of simplifications were made to limit the amount of detailed plasma information required as input to the model to facilitate the use of the model in thruster design. This simplified model gave remarkably consistant results with experimental results obtained with a given thruster geometry over substantial changes in operating conditions. The model was uncertain to about a factor of two for different thruster cathode region geometries. The design usefulness was limited by this factor of two uncertainty and by the accuracy to which the plasma parameters required as inputs to the model were specified.

  7. Multiply charged ion generation according to magnetic field configurations in Hall thruster plasmas

    NASA Astrophysics Data System (ADS)

    Kim, Holak; Lee, Seunghun; Kim, Junbum; Lim, Youbong; Choe, Wonho; KIMS Collaboration

    2016-09-01

    Plasma propulsion is the most promising techniques to operate satellites for low earth orbit as well as deep space exploration. A typical plasma propulsion system is Hall thruster (HT) that uses crossed electromagnetic fields to ionize a propellant gas and to accelerate the ionized gas. In HT the tailoring of magnetic fields is significant due to that the electron confinement in the electromagnetic fields affects thruster performances such as thrust force, specific impulse, power efficiency, and life time. We designed an anode layer HT (TAL) with the magnetic field tailoring. The TAL is possible to keep discharge in 1 2 kilovolts, which voltage is useful to obtain high specific impulse The magnetic field tailoring is adapted to minimize undesirable heat dissipations and secondary electron emissions at a wall surrounding plasma In presentation, we will report TAL performances including thrust force, specific impulse, and anode efficiency measured by a pendulum thrust stand. This mechanical measurement will be compared to the plasma diagnostics conducted by angular Faraday probe, retarding potential analyzer, and ExB probe Grant No. 2014M1A3A3A02034510.

  8. Design and Performance of 40 cm Ion Optics

    NASA Technical Reports Server (NTRS)

    Soulas, George C.

    2001-01-01

    A 40 cm ion thruster is being developed at the NASA Glenn Research Center to obtain input power and propellant throughput capabilities of 10 kW and 550 kg. respectively. The technical approach here is a continuation of the "derating" technique used for the NSTAR ion thruster. The 40 cm ion thruster presently utilizes the NSTAR ion optics aperture geometry to take advantage of the large database of lifetime and performance data already available. Dome-shaped grids were chosen for the design of the 40 cm ion optics because this design is naturally suited for large-area ion optics. Ion extraction capabilities and electron backstreaming limits for the 40 cm ion optics were estimated by utilizing NSTAR 30 cm ion optics data. A preliminary service life assessment showed that the propellant throughput goal of 550 kg of xenon may be possible with molybdenum 40 cm ion optics. One 40 cm ion optics' set has been successfully fabricated to date. Additional ion optics' sets are presently being fabricated. Preliminary performance tests were conducted on a laboratory model 40 cm ion thruster.

  9. Fifteen cm mercury ion thruster research, 1976. [performance as effected by the use of shag optics at 33 v discharge voltage

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1976-01-01

    Improvements in 15 cm diameter, SERT II, mercury ion thruster performance effected by the use of SHAG optics at 33 V discharge voltage were discussed. At a 200 eV/ion discharge power, 90 percent propellant utilization and 660 mA beam current condition a doubly-to-singly charged ion current ratio of about 4 percent was measured. Performance of the 15 cm multipole mercury thruster (optimized for length and the point of electron injection) was compared to that of divergent (SERT II) and cusped field designs and found to be comparable. The need for a magnetic baffle in the multipole thruster was identified and the preferred point of electron injection was at the upstream end of the discharge chamber. Results of preliminary tests on the effects of discharge voltage and total accelerating voltage on perveance and beam divergence characteristics of two grid ion optics were examined. Experimental data showing the effect of target temperature on sputtering rates in a mercury discharge environment were presented and a deficiency in the tests procedure was identified.

  10. Test bed ion engine development

    NASA Technical Reports Server (NTRS)

    Aston, G.; Deininger, W. D.

    1984-01-01

    A test bed ion (TBI) engine was developed to serve as a tool in exploring the limits of electrostatic ion thruster performance. A description of three key ion engine components, the decoupled extraction and amplified current (DE-AC) accelerator system, field enhanced refractory metal (FERM) hollow cathode and divergent line cusp (DLC) discharge chamber, whose designs and operating philosophies differ markedly from conventional thruster technology is given. Significant program achievements were: (1) high current density DE-AC accelerator system operation at low electric field stress with indicated feasibility of a 60 mA/sq cm argon ion beam; (2) reliable FERM cathode start up times of 1 to 2 secs. and demonstrated 35 ampere emission levels; (3) DLC discharge chamber plasma potentials negative of anode potential; and (4) identification of an efficient high plasma density engine operating mode. Using the performance projections of this program and reasonable estimates of other parameter values, a 1.0 Newton thrust ion engine is identified as a realizable technology goal. Calculations show that such an engine, comparable in beam area to a J series 30 cm thruster, could, operating on Xe or Hg, have thruster efficiencies as high as 0.76 and 0.78 respectively, with a 100 eV/ion discharge loss.

  11. Performance characteristics of the deep space 1 flight spare ion thruster long duration test, the first 21,300 hours of operation

    NASA Technical Reports Server (NTRS)

    Sengupta, A.; Anderson, J.; Brophy, J.; Rawlin, V.; Sovey, J.

    2002-01-01

    A long duration test of the DSl flight spare ion thruster (FT2) is presently being conducted at the Jet Propulsion Laboratory. To, date the thruster has accumulated over 23,500 hours of operation, and 190 kg of Xenon propellant, over 230% of the initial design life. The primary objectives of the test include the processing of 200 kg of Xenon propellant, the identification of unknown failure modes, the characterization and drivers of these failure modes, and to measure performance degradation as the thruster wears. The test is fitted with an extensive array of diagnostics to measure engine wear and performance degradation. To date the most notable erosion processes include severe discharge cathode keeper erosion, accelerator grid erosion, reduction in electrical isolation of the neutralizer assembly, and deposit formation within the neutralizer orifice, reducing margin from plume mode. Over the past 23,500 hours of operation, performance degradation has been minimal, and it is anticipated that the above erosion processes will not preclude the thruster from processing over 200 kg of Xenon.

  12. Adapting magnetoelectrostatic containment to inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Ramsey, W. D.; James, E. L.

    1981-01-01

    Two different types of 12 cm magnetoelectrostatic containment (MESC) ion thrusters have been adapted to argon-xenon operation. Discharge chamber optimization produced excellent performance with both the hexagonal and hemispherical shaped thrusters. The hemispherical thruster design yielded the best performance, ionizing 75 to 96 percent of the xenon propellant with a discharge energy consumption rate of 185 to 320 eV/ion. Argon operation of the same thruster achieved 60 to 80 percent propellant ionization at 215 to 370 eV/ion.

  13. Very high delta-V missions to the edge of the solar system and beyond enabled by the dual-stage 4-grid ion thruster concept

    NASA Astrophysics Data System (ADS)

    Bramanti, C.; Izzo, D.; Samaraee, T.; Walker, R.; Fearn, D.

    2009-04-01

    A new and innovative type of gridded ion thruster, the "Dual-Stage 4-Grid" or DS4G concept, has been proposed and its predicted high performance validated under an ESA research, development and test programme. The DS4G concept is able to operate at very high specific impulse and thrust density values well in excess of conventional 3-grid ion thrusters at the expense of a higher power-to-thrust ratio. This makes it a possible candidate for ambitious missions requiring very high delta-V capability and high power. Such missions include 100 kW-level multi-ton probes based on nuclear and solar electric propulsion (SEP) to distant Kuiper Belt Object and inner Oort cloud objects, and to the Local Interstellar medium. In this paper, the DS4G concept is introduced and its application to this mission class is investigated. Benefits of using the DS4G over conventional thrusters include reduced transfer time and increased payload mass, if suitably advanced lightweight power system technologies are developed. A mission-level optimisation is performed (launch, spacecraft system design and low-thrust trajectory combined) in order to find design solutions with minimum transfer time, maximum scientific payload mass, and to explore the influence of power system specific mass. It is found that the DS4G enables an 8-ton spacecraft with a payload mass of 400 kg, equipped with a 65 kW nuclear reactor with specific mass 25 kg/kW (e.g. Topaz-type with Brayton cycle conversion) to reach 200 AU in 23 years after an Earth escape launch by Ariane 5. In this scenario, the optimum specific impulse for the mission is over 10,000 s, which is well within the capabilities of a single 65 kW DS4G thruster. It is also found that an interstellar probe mission to 200 AU could be accomplished in 25 years using a "medium-term" SEP system with a lightweight 155 kW solar array (2 kg/kW specific mass) and thruster PPU (3.7 kg/kW) and an Earth escape launch on Ariane 5. In this case, the optimum specific

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

    SciTech Connect

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

    2012-11-15

    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.

  15. Modeling of life limiting phenomena in the discharge chamber of an electron bombardment ion thruster

    NASA Technical Reports Server (NTRS)

    Handoo, Arvind K.; Ray, Pradosh K.

    1991-01-01

    An experimental facility to study the low energy sputtering of metal surfaces with ions produced by an ion gun is described. The energy of the ions ranged from 10 to 500 eV. Cesium ions with energies from 100 to 500 eV were used initially to characterize the operation of the ion gun. Next, argon and xenon ions were used to measure the sputtering yields of cobalt (Co), Cadmium (Cd), and Chromium (Cr) at an operating temperature of 2x10(exp -5) Torr. The ion current ranged from 0.0135 micro-A at 500 eV. The targets were electroplated on a copper substrate. The surface density of the electroplated material was approx. 50 micro-g/sq cm. The sputtered atoms were collected on an aluminum foil surrounding the target. Radioactive tracers were used to measure the sputtering yields. The sputtering yields of Cr were found to be much higher than those of Co and Cd. The yields of Co and Cd were comparable, with Co providing the higher yields. Co and Cd targets were observed to sputter at energies as low as 10 eV for both argon and xenon ions. The Cr yields could not be measured below 20 eV for argon ions and 15 eV for xenon ions. On a linear scale the yield energy curves near the threshold energies exhibit a concave nature.

  16. Improving the Total Impulse Capability of the NSTAR Ion Thruster With Thick-Accelerator-Grid Ion Optics

    NASA Technical Reports Server (NTRS)

    Soulas, George C.

    2001-01-01

    The results of performance tests with thick-accelerator-grid (TAG) ion optics are presented. TAG ion optics utilize a 50 percent thicker accelerator grid to double ion optics' service life. NSTAR ion optics were also tested to provide a baseline performance for comparison. Impingement-limited total voltages for the TAG ion optics were only 0 to 15 V higher than those of the NSTAR ion optics. Electron backstreaming limits for the TAG ion optics were 3 to 9 V higher than those for the NSTAR optics due to the increased accelerator grid thickness for the TAG ion optics. Screen grid ion transparencies for the TAG ion optics were only about 2 percent lower than those for the NSTAR optics, reflecting the lower physical screen grid open area fraction of the TAG ion optics. Accelerator currents for the TAG ion optics were 19 to 43 percent greater than those for the NSTAR ion optics due, in part, to a sudden increase in accelerator current during TAG ion optics' performance tests for unknown reasons and to the lower-than-nominal accelerator aperture diameters. Beam divergence half-angles that enclosed 95 percent of the total beam current and beam divergence thrust correction factors for the TAG ion optics were within 2 degrees and 1 percent, respectively, of those for the NSTAR ion optics.

  17. Radiation hardness of 30 cm long CsI(Tl) crystals

    NASA Astrophysics Data System (ADS)

    Longo, S.; Roney, J. M.

    2016-08-01

    Measurements of the degradation in performance of 30 cm long CsI(Tl) scintillation crystals exposed to 1 MeV photon doses of 2, 10, 35, 100 and 1000 Gy are presented. The light yield, light yield longitudinal non-uniformity, scintillation decay times, energy resolution and timing resolution of a set of spare crystals from the BABAR and Belle experiments are studied as a function of these doses. In addition, a model that describes the plateau observed in the light output loss as a function of dose in terms of increase in concentrations of absorption centres with irradiation is presented.

  18. Improvement of the Power Control Unit for Ion Thruster to Cope with Milli-Newton Range RIT

    NASA Astrophysics Data System (ADS)

    Ceruti, Luca; Polli, Aldo; Galantini, Paolo

    2014-08-01

    The recent development and testing activities of a miniaturized Radio-Frequency Ion Thruster, with relevant ancillary elements, in the range of 10 to 100 micro-Newtons, joined with past flight heritage in the milli-Newton range (RIT-10 for Artemis), shows an appealing capability of such an electrical propulsion technology to support thrust in a wide range of space applications from very fine attitude control up to deorbiting of small-medium satellites. As expectable, this implies that the mentioned ancillary elements (mainly Radio-Frequency Generator and Power Control Unit) require adaptation to the different requirements imposed to different missions and thrust ranges. Regarding the Power Control Unit different power levels, both the controllability requirements and the spacecraft interfaces impose non negligible adaptation leading to significant increase of development activities and associated cost (nonrecurring) increase. From that and with the main purpose to minimize such impacts and provide reliable equipments, Selex ES since a few years is devoting maximum attention in the incremental innovation of the existing design in order to maximize their reuse.

  19. The Impact of Back-Sputtered Carbon on the Accelerator Grid Wear Rates of the NEXT and NSTAR Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Soulas, George C.

    2013-01-01

    A study was conducted to quantify the impact of back-sputtered carbon on the downstream accelerator grid erosion rates of the NASA's Evolutionary Xenon Thruster (NEXT) Long Duration Test (LDT1). A similar analysis that was conducted for the NASA's Solar Electric Propulsion Technology Applications Readiness Program (NSTAR) Life Demonstration Test (LDT2) was used as a foundation for the analysis developed herein. A new carbon surface coverage model was developed that accounted for multiple carbon adlayers before complete surface coverage is achieved. The resulting model requires knowledge of more model inputs, so they were conservatively estimated using the results of past thin film sputtering studies and particle reflection predictions. In addition, accelerator current densities across the grid were rigorously determined using an ion optics code to determine accelerator current distributions and an algorithm to determine beam current densities along a grid using downstream measurements. The improved analysis was applied to the NSTAR test results for evaluation. The improved analysis demonstrated that the impact of back-sputtered carbon on pit and groove wear rate for the NSTAR LDT2 was negligible throughout most of eroded grid radius. The improved analysis also predicted the accelerator current density for transition from net erosion to net deposition considerably more accurately than the original analysis. The improved analysis was used to estimate the impact of back-sputtered carbon on the accelerator grid pit and groove wear rate of the NEXT Long Duration Test (LDT1). Unlike the NSTAR analysis, the NEXT analysis was more challenging because the thruster was operated for extended durations at various operating conditions and was unavailable for measurements because the test is ongoing. As a result, the NEXT LDT1 estimates presented herein are considered preliminary until the results of future post-test analyses are incorporated. The worst-case impact of carbon

  20. The Impact of Back-Sputtered Carbon on the Accelerator Grid Wear Rates of the NEXT and NSTAR Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Soulas, George C.

    2013-01-01

    A study was conducted to quantify the impact of back-sputtered carbon on the downstream accelerator grid erosion rates of the NEXT (NASA's Evolutionary Xenon Thruster) Long Duration Test (LDT1). A similar analysis that was conducted for the NSTAR (NASA's Solar Electric Propulsion Technology Applications Readiness Program) Life Demonstration Test (LDT2) was used as a foundation for the analysis developed herein. A new carbon surface coverage model was developed that accounted for multiple carbon adlayers before complete surface coverage is achieved. The resulting model requires knowledge of more model inputs, so they were conservatively estimated using the results of past thin film sputtering studies and particle reflection predictions. In addition, accelerator current densities across the grid were rigorously determined using an ion optics code to determine accelerator current distributions and an algorithm to determine beam current densities along a grid using downstream measurements. The improved analysis was applied to the NSTAR test results for evaluation. The improved analysis demonstrated that the impact of back-sputtered carbon on pit and groove wear rate for the NSTAR LDT2 was negligible throughout most of eroded grid radius. The improved analysis also predicted the accelerator current density for transition from net erosion to net deposition considerably more accurately than the original analysis. The improved analysis was used to estimate the impact of back-sputtered carbon on the accelerator grid pit and groove wear rate of the NEXT Long Duration Test (LDT1). Unlike the NSTAR analysis, the NEXT analysis was more challenging because the thruster was operated for extended durations at various operating conditions and was unavailable for measurements because the test is ongoing. As a result, the NEXT LDT1 estimates presented herein are considered preliminary until the results of future posttest analyses are incorporated. The worst-case impact of carbon back

  1. Ion Heating Experiments in a Supersonic Plasma Flow for an Advanced Plasma Thruster

    NASA Astrophysics Data System (ADS)

    Ando, Akira; Hosokawa, Yohei; Hatanaka, Motoi; Yagai, Tsuyoshi; Tobari, Hiroyuki; Hattori, Kunihiko; Inutake, Masaaki

    2003-10-01

    In the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) project in NASA, the combined system of the ion cyclotron heating and the magnetic nozzle is proposed to control a ratio od specific impulse to thrust at constant power. By now, few attempt of a direct ion heating for fast flowing plasma by waves has been done. Ion heating in a fast flowing plasma might be difficult because of the short transit time for ions to pass through a heating region only once and the modification of ion cyclotron resonance due to the effect of Doppler shift. Ion heating experiments are performed in a fast flowing plasma produced by Magneto-Plasma-Dynamic Arcjet (MPDA) operated with an externally-applied magnetic field. RF waves with an ion cyclotron range of frequency is excited by a pair of loop antennas or a helical antenna. An increase of plasma stored energy measured by a diamagnetic loop coil is observed when the waves are excited with various azimuthal mode numbers in several magnetic nozzle configurations. It is most effective to heat ions to excite the waves with an azimuthal mode number of m=±1. Dispersion relations of the propagating wave are obtained and compared with theoretical ones.

  2. A Comparison of Ion Acceleration Characteristics for Krypton and Xenon Propellants within a 600 Watt Hall Thruster

    DTIC Science & Technology

    2012-07-20

    fluctuation.8,9 However, care must be taken to ensure that the relative effects of these phenomena are separable. In addition, magnetic ( Zeeman effect ...capability. This work compares the internal propellant acceleration of krypton within a laboratory medium power Hall effect thruster to historical xenon...Watt Hall Effect Thruster William A. Hargus, Jr.∗ Gregory M. Azarnia† Michael R. Nakles‡ Air Force Research Laboratory, Edwards Air Force Base, CA

  3. Theoretical investigations on plasma processes in the Kaufman thruster. [electron and ion velocity distribution

    NASA Technical Reports Server (NTRS)

    Wilhelm, H. E.

    1974-01-01

    An analysis of the sputtering of metal surfaces and grids by ions of medium energies is given and it is shown that an exact, nonlinear, hyperbolic wave equation for the temperature field describes the transient transport of heat in metals. Quantum statistical and perturbation theoretical analysis of surface sputtering by low energy ions are used to develop the same expression for the sputtering rate. A transport model is formulated for the deposition of sputtered atoms on system components. Theoretical efforts in determining the potential distribution and the particle velocity distributions in low pressure discharges are briefly discussed.

  4. Helical plasma thruster

    SciTech Connect

    Beklemishev, A. D.

    2015-10-15

    A new scheme of plasma thruster is proposed. It is based on axial acceleration of rotating magnetized plasmas in magnetic field with helical corrugation. The idea is that the propellant ionization zone can be placed into the local magnetic well, so that initially the ions are trapped. The E × B rotation is provided by an applied radial electric field that makes the setup similar to a magnetron discharge. Then, from the rotating plasma viewpoint, the magnetic wells of the helically corrugated field look like axially moving mirror traps. Specific shaping of the corrugation can allow continuous acceleration of trapped plasma ions along the magnetic field by diamagnetic forces. The accelerated propellant is expelled through the expanding field of magnetic nozzle. By features of the acceleration principle, the helical plasma thruster may operate at high energy densities but requires a rather high axial magnetic field, which places it in the same class as the VASIMR{sup ®} rocket engine.

  5. Helical plasma thruster

    NASA Astrophysics Data System (ADS)

    Beklemishev, A. D.

    2015-10-01

    A new scheme of plasma thruster is proposed. It is based on axial acceleration of rotating magnetized plasmas in magnetic field with helical corrugation. The idea is that the propellant ionization zone can be placed into the local magnetic well, so that initially the ions are trapped. The E × B rotation is provided by an applied radial electric field that makes the setup similar to a magnetron discharge. Then, from the rotating plasma viewpoint, the magnetic wells of the helically corrugated field look like axially moving mirror traps. Specific shaping of the corrugation can allow continuous acceleration of trapped plasma ions along the magnetic field by diamagnetic forces. The accelerated propellant is expelled through the expanding field of magnetic nozzle. By features of the acceleration principle, the helical plasma thruster may operate at high energy densities but requires a rather high axial magnetic field, which places it in the same class as the VASIMR® rocket engine.

  6. Plasma-Surface Interactions in Electric Thrusters

    NASA Astrophysics Data System (ADS)

    Goebel, Dan

    2013-09-01

    Of critical importance in electric propulsion missions in space is thruster life, which is determined to a large extent by wall erosion from plasma-materials interactions. While the plasmas generated in different thrusters vary, the particle fluxes, energies and temperatures in contact with the walls are somewhat similar. The erosion rates are then determined by details of materials, incident angles, etc. In ion and Hall thrusters commonly used today, for example, cathode life is determined by low energy (<=100 eV) Xe ion erosion of the cathode electrodes. Erosion of ion thruster accelerator grids is dominated by charge exchange ion bombardment with energies of 200 to 400 V. The incident angle of these ions is near normal, but the sputtered particles are ejected with a butterfly distribution that directs particles along the thruster axis and causes build up of material on the upstream and downstream surfaces. In Hall thrusters, the plasma materials interactions at the wall are complicated because the walls are typically ceramic and selected for a low secondary electron yield for thruster performance. The erosion rates at the wall vary due to non-uniform plasma contact with the surface causing grooves and surface changes. These effects will be discussed for various thrusters.

  7. Inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1979-01-01

    Inert gas thrusters considered for space propulsion systems were investigated. Electron diffusion across a magnetic field was examined utilizing a basic model. The production of doubly charged ions was correlated using only overall performance parameters. The use of this correlation is therefore possible in the design stage of large gas thrusters, where detailed plasma properties are not available. Argon hollow cathode performance was investigated over a range of emission currents, with the positions of the inert, keeper, and anode varied. A general trend observed was that the maximum ratio of emission to flow rate increased at higher propellant flow rates. It was also found that an enclosed keeper enhances maximum cathode emission at high flow rates. The maximum cathode emission at a given flow rate was associated with a noisy high voltage mode. Although this mode has some similarities to the plume mode found at low flows and emissions, it is encountered by being initially in the spot mode and increasing emission. A detailed analysis of large, inert-gas thruster performance was carried out. For maximum thruster efficiency, the optimum beam diameter increases from less than a meter at under 2000 sec specific impulse to several meters at 10,000 sec. The corresponding range in input power ranges from several kilowatts to megawatts.

  8. Influence of Triply-Charged Ions and Ionization Cross-Sections in a Hybrid-PIC Model of a Hall Thruster Discharge

    NASA Technical Reports Server (NTRS)

    Smith, Brandon D.; Boyd, Iain D.; Kamhawi, Hani

    2014-01-01

    The sensitivity of xenon ionization rates to collision cross-sections is studied within the framework of a hybrid-PIC model of a Hall thruster discharge. A revised curve fit based on the Drawin form is proposed and is shown to better reproduce the measured crosssections at high electron energies, with differences in the integrated rate coefficients being on the order of 10% for electron temperatures between 20 eV and 30 eV. The revised fit is implemented into HPHall and the updated model is used to simulate NASA's HiVHAc EDU2 Hall thruster at discharge voltages of 300, 400, and 500 V. For all three operating points, the revised cross-sections result in an increase in the predicted thrust and anode efficiency, reducing the error relative to experimental performance measurements. Electron temperature and ionization reaction rates are shown to follow the trends expected based on the integrated rate coefficients. The effects of triply-charged xenon are also assessed. The predicted thruster performance is found to have little or no dependence on the presence of triply-charged ions. The fraction of ion current carried by triply-charged ions is found to be on the order of 1% and increases slightly with increasing discharge voltage. The reaction rates for the 0?III, I?III, and II?III ionization reactions are found to be of similar order of magnitude and are about one order of magnitude smaller than the rate of 0?II ionization in the discharge channel.

  9. Constrained sheath optics for high thrust density, low specific impulse ion thrusters

    NASA Technical Reports Server (NTRS)

    Wilbur, Paul J.; Han, Jian-Zhang

    1987-01-01

    The results of an experimental study showing that a contoured, fine wire mesh attached to the screen grid can be used to control the divergence characteristics of ion beamlets produced at low net-to-total accelerating voltage ratios are presented. The influence of free and constrained-sheath optics systems on beamlet divergence characteristics are found to be similar in the operating regime investigated, but it was found that constrained-sheath optics systems can be operated at higher perveance levels than free-sheath ones. The concept of a fine wire interference probe that can be used to study ion beamlet focusing behavior is introduced. This probe is used to demonstrate beamlet focusing to a diameter about one hundreth of the screen grid extraction aperture diameter. Additional testing is suggested to define an optimally contoured mesh that could yield well focused beamlets at net-to-total accelerating voltage ratios below about 0.1.

  10. Design and Installation of a Field Ionization Test Chamber for Ion Thrusters

    DTIC Science & Technology

    2011-12-01

    Romano Second Reader: Oscar Biblarz THIS PAGE INTENTIONALLY LEFT BLANK i REPORT DOCUMENTATION PAGE Form Approved OMB No. 0704–0188 Public...Paul W. Camp Approved by: Sebastian Osswald Thesis Advisor Marcello Romano Thesis Co-Advisor Oscar Biblarz Second Reader Knox T...in its success. I would also like to thank Professor Oscar Biblarz, for his wealth of knowledge on ion propulsion he imparted to me. A special

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

    DTIC Science & Technology

    2016-09-14

    electric transfer vehicles that would strain world-wide xenon production. This work compares the internal propellant acceleration of krypton ions...acceleration rate is lower and produces a lower effective electric field. As a result, energy conversion is lower than xenon for this flow matched case. In...lower cost replacement for xenon, may optimize to similar or potentially higher performance, and is enabling for very large solar electric transfer

  12. Cylindrical Hall Thrusters with Permanent Magnets

    SciTech Connect

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

    2010-10-18

    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. __________________________________________________

  13. Cylindrical Hall thrusters with permanent magnets

    NASA Astrophysics Data System (ADS)

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

    2010-11-01

    The use of permanent magnets instead of electromagnet coils for low power Hall thrusters can offer a significant reduction in 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 50-300 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.

  14. How to Directly Image a Habitable Planet Around Alpha Centauri with a 30cm Space Telescope.

    NASA Astrophysics Data System (ADS)

    Belikov, R.; Bendek, E.; Thomas, S.; Black, D.

    2014-12-01

    More than 1,700 exoplanets have been discovered to date, including a handful of potentially habitable ones. There is on average more than one planet per star, and estimates of occurrence rates for potentially habitable planets (eta_Earth) from the Kepler mission range between 5 and 50%. Several mission concepts have been studied to directly image planets around nearby stars. Direct imaging enables spectroscopic detection of biomarkers such as atmospheric oxygen and methane, which would be highly suggestive of extraterrestrial life. It is commonly thought that directly imaging a potentially habitable exoplanet requires telescopes with apertures of at least 1m, costing at least $1B, and launching no earlier than the 2020s. A notable exception to this is Alpha Centauri. The system contains two Sun-like stars with a wide separation that allows dynamically stable habitable zones around either star. Habitable zones span about 0.5-1" in stellocentric angle, 3x wider than around any other FGKM star. A 30cm visible light space telescope is sufficient to resolve the habitable zone and detect a potentially habitable planet in minutes with ideal components, or days with realistic ones. We are developing a mission concept called ACEND (Alpha Centauri Direct Imager) consisting of a 30cm primary, a Phase-Induced Amplitude Apodization coronagraph, and a wavefront control system. It is designed to suppress the light leak from both stars and directly image their planetary systems in 3 color channels, including the capability to detect potentially habitable planets. Color imaging is sufficient to differentiate Venus-like, Earth-like, and Mars-like planets from each other and establish the presence of Earth-pressure atmosphere through Rayleigh scattering. Two factors make it possible to realize the requirements of ACEND (most notably 10^10 contrast) on a small budget and fast schedule: (a) ACEND will collect a long continuous sequence of images on Alpha Centauri A and B for 2 years

  15. Electron dynamics in Hall thruster

    NASA Astrophysics Data System (ADS)

    Marini, Samuel; Pakter, Renato

    2015-11-01

    Hall thrusters are plasma engines those use an electromagnetic fields combination to confine electrons, generate and accelerate ions. Widely used by aerospace industries those thrusters stand out for its simple geometry, high specific impulse and low demand for electric power. Propulsion generated by those systems is due to acceleration of ions produced in an acceleration channel. The ions are generated by collision of electrons with propellant gas atoms. In this context, we can realize how important is characterizing the electronic dynamics. Using Hamiltonian formalism, we derive the electron motion equation in a simplified electromagnetic fields configuration observed in hall thrusters. We found conditions those must be satisfied by electromagnetic fields to have electronic confinement in acceleration channel. We present configurations of electromagnetic fields those maximize propellant gas ionization and thus make propulsion more efficient. This work was supported by CNPq.

  16. Power processing technology for spacecraft primary ion propulsion

    NASA Technical Reports Server (NTRS)

    Biess, J. J.; Inouye, L. Y.; Frye, R. J.

    1980-01-01

    Advanced technologies developed in support of Ion Propulsion power processing, including the power circuitry portion of the Series L-C Resonant Inverter, Beam Supply, power components, packaging and heat pipe cooling of the 30 cm Ion Engine Power Processor are described. Both the transistorized and SCR versions of the Series L-C Resonant Inverter Beam Supply are discussed. A BIMOD Ion Thruster/Power Processor Prototype Assembly is undergoing environmental and life testing. These advanced technologies can be applied advantageously to other applications of future high power space power processing equipment.

  17. Long Term Multiplication Behavior Studies of the 30cmx 30cm prototype Gas electron Multiplier

    NASA Astrophysics Data System (ADS)

    Ng, Ying Wun Yvonne; Yu, Jaehoon; Park, Seongtae; White, Andy

    2014-03-01

    The Gas Electron Multiplier (GEM) technology is one of the next generation radiation detector technologies that utilized the ionization in gaseous medium and the electron avalanche to detect a magnified charge value from various radiation and charge particles. With its low building cost, low discharge rate and high resolution, GEM is currently being considered to be one of the candidate gap detectors for the International Linear Collider (ILC) in Japan. It is therefore of crucial for us to study the long term stability of amplification power of the detector. Using cosmic radiation as our radiation source, data has been taken continuously in the past 2 years by the high energy physics group in University of Texas at Arlington to characterize the stability of the 30cmx30cm detector. Effect of atmospheric pressure to the detector amplification is eliminated by a correction algorithm. Noise study has been done to eliminate excessive noise produced by the detector as well as its readout chip. Result shows that the detector gives us a stable 35fC average MPV for the cosmic MIPs with few fC of chamber noise and about 0.5 of chip noise. GEM should work well as a digital calorimeter for uses in the ILC project.

  18. The 30 cm radio flux as a solar proxy for thermosphere density modelling

    NASA Astrophysics Data System (ADS)

    Dudok de Wit, Thierry; Bruinsma, Sean

    2017-03-01

    The 10.7 cm radio flux (F10.7) is widely used as a proxy for solar UV forcing of the upper atmosphere. However, radio emissions at other centimetric wavelengths have been routinely monitored since the 1950 s, thereby offering prospects for building proxies that may be better tailored to space weather needs. Here we advocate the 30 cm flux (F30) as a proxy that is more sensitive than F10.7 to longer wavelengths in the UV and show that it improves the response of the thermospheric density to solar forcing, as modelled with DTM (Drag Temperature Model). In particular, the model bias drops on average by 0-20% when replacing F10.7 by F30; it is also more stable (the standard deviation of the bias is 15-40% smaller) and the density variation at the the solar rotation period is reproduced with a 35-50% smaller error. We compare F30 to other solar proxies and discuss its assets and limitations.

  19. Development Status of a Power Processing Unit for Low Power Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Pinero, Luis R.; Bowers, Glen E.; Lafontaine, Eric M.

    2000-01-01

    An advanced breadboard Power Processing Unit (PPU) for a low power ion propulsion system incorporating mass reduction techniques was designed and fabricated. As a result of similar output current requirements, the discharge supply was also used to provide the neutralizer heater and discharge heater functions by using three relays to switch the output connections. This multi-function supply reduces to four the number of power converters needed to produce the required six electrical outputs. Switching frequencies of 20 and 50 kHz were chosen as a compromise between the size of the magnetic components and switching losses. The advanced breadboard PPU is capable of a maximum total output power of 0.47 kW. Its component mass is 0.65 kg and its total mass 1.9 kg. The total efficiency at full power is 0.89.

  20. Operating characteristics of a hollow-cathode neutralizer for 5 and 8 centimeter-diameter electron bombardment mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Weigand, A. J.

    1975-01-01

    Thin-tip 0.3-cm-outside-diameter hollow-cathode neutralizers were used to investigate causes of neutralizer tip erosion experienced in thruster endurance tests. Bell-jar tests indicated that neutralizers with new rolled tantalum foil inserts coated with an emissive mixture eroded very little over the neutral flow rates investigated (3 to 10 mA) for simulated 5- and 8-cm-diameter thruster neutralizer conditions. Tip erosion rates of neutralizers operated with no insert or emissive mixture increased by two orders of magnitude for both configurations as the neutral flow rate decreased. Spectroscopic analysis of the discharge plasma from neutralizers operated with inserts coated with the emissive mixture detected tungsten at all neutral flow rates for both thruster neutralizer conditions. The only source of tungsten was the tip. Therefore, detection of tungsten indicated neutralizer tip erosion. Barium, an element of the emissive mixture, was detected at low neutral flow rates for the 5-cm-diameter thruster neutralizer operating condition only.

  1. Electric thruster research

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1981-01-01

    The multipole discharge chamber of an electrostatic ion thruster is discussed. No reductions in discharge losses were obtained, despite repeated demonstration of anode potentials more positive than the bulk of the discharge plasma. The penalty associated with biased anode operation was reduced as the magnetic integral above the biased anodes was increased. The hollow cathode is discussed. The experimental configuration of the Hall current thruster had a uniform field throughout the ion generation and acceleration regions. To obtain reliable ion generation, it was necessary to reduce the magnetic field strength, to the point where excessive electron backflow was required to establish ion acceleration. The theoretical study of ion acceleration with closed electron drift paths resulted in two classes of solutions. One class has the continuous potential variation in the acceleration region that is normally associated with a Hall current accelerator. The other class has an almost discontinuous potential step near the anode end of the acceleration region. This step includes a significant fraction of the total acceleration potential difference.

  2. Performance Evaluation of 40 cm Ion Optics for the NEXT Ion Engine

    NASA Technical Reports Server (NTRS)

    Soulas, George C.; Haag, Thomas W.; Patterson, Michael J.

    2002-01-01

    The results of performance tests with two 40 cm ion optics sets are presented and compared to those of 30 cm ion optics with similar aperture geometries. The 40 cm ion optics utilized both NSTAR and TAG (Thick-Accelerator-Grid) aperture geometries. All 40 cm ion optics tests were conducted on a NEXT (NASA's Evolutionary Xenon Thruster) laboratory model ion engine. Ion optics performance tests were conducted over a beam current range of 1.20 to 3.52 A and an engine input power range of 1.1 to 6.9 kW. Measured ion optics' performance parameters included near-field radial beam current density profiles, impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, beam divergence angles, and start-up transients. Impingement-limited total voltages for 40 cm ion optics with the NSTAR aperture geometry were 60 to 90 V lower than those with the TAG aperture geometry. This difference was speculated to be due to an incomplete burn-in of the TAG ion optics. Electron backstreaming limits for the 40 cm ion optics with the TAG aperture geometry were 8 to 19 V higher than those with the NSTAR aperture geometry due to the thicker accelerator grid of the TAG geometry. Because the NEXT ion engine provided beam flatness parameters that were 40 to 63 percent higher than those of the NSTAR ion engine, the 40 cm ion optics outperformed the 30 cm ion optics.

  3. Stationary Plasma Thruster Plume Characteristics

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Manzella, David H.

    1994-01-01

    Stationary Plasma Thrusters (SPT's) are being investigated for application to a variety of near-term missions. This paper presents the results of a preliminary study of the thruster plume characteristics which are needed to assess spacecraft integration requirements. Langmuir probes, planar probes, Faraday cups, and a retarding potential analyzer were used to measure plume properties. For the design operating voltage of 300 V the centerline electron density was found to decrease from approximately 1.8 x 10 exp 17 cubic meters at a distance of 0.3 m to 1.8 X 10 exp 14 cubic meters at a distance of 4 m from the thruster. The electron temperature over the same region was between 1.7 and 3.5 eV. Ion current density measurements showed that the plume was sharply peaked, dropping by a factor of 2.6 within 22 degrees of centerline. The ion energy 4 m from the thruster and 15 degrees off-centerline was approximately 270 V. The thruster cathode flow rate and facility pressure were found to strongly affect the plume properties. In addition to the plume measurements, the data from the various probe types were used to assess the impact of probe design criteria

  4. How to Directly Image a Habitable Planet Around Alpha Centauri with a ~30cm Space Telescope

    NASA Astrophysics Data System (ADS)

    Belikov, Ruslan; Acend Team, Acesat Team

    2015-01-01

    Several mission concepts are being studied to directly image planets around nearby stars. Direct imaging enables spectroscopic detection of biomarkers such as atmospheric oxygen and methane, which would be highly suggestive of extraterrestrial life. It is commonly thought that directly imaging a potentially habitable exoplanet requires telescopes with apertures of at least 1m, costing at least 1B, and launching no earlier than the 2020s.A notable exception to this is Alpha Centauri (A and B), which is an extreme outlier among FGKM stars in terms of apparent habitable zone size. Specifically, Alpha Centauri habitable zones span about 0.5-1' in stellocentric angle, ~3x wider than around any other FGKM star. This enables a ~30cm visible light space telescope equipped with a modern high performance coronagraph or starshade to resolve the habitable zone at high contrast and directly image any potentially habitable planet that may exist in the system. Due to the extreme apparent brightness of the stars, exposure times can be as short as minutes with ideal components, or days with realistic ones. This makes it possible to do color photometry on potentially habitable planets sufficient to differentiate Venus-like, Earth-like, and Mars-like planets from each other and establish the presence of Earth-pressure atmosphere through Rayleigh scattering.The raw contrast requirements for such an instrument can be relaxed to 1e-8 if the mission spends 2 years collecting tens of thousands of images on the same target, enabling a factor of 500-1000 speckle suppression in post processing. The light leak from both stars is controllable with a special wavefront control algorithm known as Multi-Star Wavefront Control (MSWC), which independently suppresses diffraction and aberrations from both stars using independent modes on the deformable mirror (see Thomas et al. at this conference).The presentation will describe the general studies and calculations in more detail and briefly present

  5. Pulsed Plasma Thruster Contamination

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Arrington, Lynn A.; Pencil, Eric J.; Carter, Justin; Heminger, Jason; Gatsonis, Nicolas

    1996-01-01

    Pulsed Plasma Thrusters (PPT's) are currently baselined for the Air Force Mightysat II.1 flight in 1999 and are under consideration for a number of other missions for primary propulsion, precision positioning, and attitude control functions. In this work, PPT plumes were characterized to assess their contamination characteristics. Diagnostics included planar and cylindrical Langmuir probes and a large number of collimated quartz contamination sensors. Measurements were made using a LES 8/9 flight PPT at 0.24, 0.39, 0.55, and 1.2 m from the thruster, as well as in the backflow region behind the thruster. Plasma measurements revealed a peak centerline ion density and velocity of approx. 6 x 10(exp 12) cm(exp -3) and 42,000 m/s, respectively. Optical transmittance measurements of the quartz sensors after 2 x 10(exp 5) pulses showed a rapid decrease in plume contamination with increasing angle from the plume axis, with a barely measurable transmittance decrease in the ultraviolet at 90 deg. No change in optical properties was detected for sensors in the backflow region.

  6. Cathode Effects in Cylindrical Hall Thrusters

    SciTech Connect

    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.

  7. Propulsion Instruments for Small Hall Thruster Integration

    NASA Technical Reports Server (NTRS)

    Johnson, Lee K.; Conroy, David G.; Spanjers, Greg G.; Bromaghim, Daron R.

    2001-01-01

    Planning and development are underway for the propulsion instrumentation necessary for the next AFRL electric propulsion flight project, which includes both a small Hall thruster and a micro-PPT. These instruments characterize the environment induced by the thruster and the associated data constitute part of a 'user's manual' for these thrusters. Several instruments probe the back-flow region of the thruster plume, and the data are intended for comparison with detailed numerical models in this region. Specifically, an ion probe is under development to determine the energy and species distributions, and a Langmuir probe will be employed to characterize the electron density and temperature. Other instruments directly measure the effects of thruster operation on spacecraft thermal control surfaces, optical surfaces, and solar arrays. Specifically, radiometric, photometric, and solar-cell-based sensors are under development. Prototype test data for most sensors should be available, together with details of the instrumentation subsystem and spacecraft interface.

  8. A Preliminary Investigation of Hall Thruster Technology

    NASA Technical Reports Server (NTRS)

    Gallimore, Alec D.

    1997-01-01

    A three-year, NASA/BMDO-sponsored experimental program to conduct performance and plume plasma property measurements on two Russian Stationary Plasma Thrusters (SPTs) has been completed. The program utilized experimental facilitates at the University of Michigan's Plasmadynamics and Electric Propulsion Laboratory (PEPL). The main features of the proposed effort were as follows: We Characterized Hall thruster [and arcjet] performance by measuring ion exhaust velocity with probes at various thruster conditions. Used a variety of probe diagnostics in the thruster plume to measure plasma properties and flow properties including T(sub e) and n(sub e), ion current density and ion energy distribution, and electric fields by mapping plasma potential. Used emission spectroscopy to identify species within the plume and to measure electron temperatures.

  9. NEXT Thruster Component Verification Testing

    NASA Technical Reports Server (NTRS)

    Pinero, Luis R.; Sovey, James S.

    2007-01-01

    Component testing is a critical part of thruster life validation activities under NASA s Evolutionary Xenon Thruster (NEXT) project testing. The high voltage propellant isolators were selected for design verification testing. Even though they are based on a heritage design, design changes were made because the isolators will be operated under different environmental conditions including temperature, voltage, and pressure. The life test of two NEXT isolators was therefore initiated and has accumulated more than 10,000 hr of operation. Measurements to date indicate only a negligibly small increase in leakage current. The cathode heaters were also selected for verification testing. The technology to fabricate these heaters, developed for the International Space Station plasma contactor hollow cathode assembly, was transferred to Aerojet for the fabrication of the NEXT prototype model ion thrusters. Testing the contractor-fabricated heaters is necessary to validate fabrication processes for high reliability heaters. This paper documents the status of the propellant isolator and cathode heater tests.

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

    NASA Technical Reports Server (NTRS)

    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.

  11. Direct Drive for Low Power Hall Thrusters

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.

    2005-01-01

    Due to recent studies, NASA has initiated the development of a low power Hall thruster for discovery class missions. The potential advantages of a low power Hall thruster is primarily due to its high efficiency operation at low power and its lower complexity compared to ion engines. Direct drive is another method of reducing the complexity of a Hall thruster system while improving its efficiency. The technical challenges associated with this technology are reported. Additionally, the benefits of this technology are discussed based on parametric studies and mission analysis.

  12. Performance of large inert-gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1981-01-01

    The performance of large inert-gas thrusters is predicted based on present knowledge of ion optics performance and discharge chamber operation. Calculated performance data are given for argon and xenon propellants. The effect of varying propellant utilization and thruster diameter is discussed and the optimum choice of beam diameter for very large systems is indicated for low, intermediate, and high specific impulses. Optimum discharge chamber depths are also specified. Although detailed design considerations may modify the predictions, the general trends indicated should still be useful for directing future technology efforts and evaluating mission studies involving large thrusters.

  13. Plasma processes in inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1979-01-01

    Inert gas thrusters, particularly with large diameters, have continued to be of interest for space propulsion applications. Two plasma processes are treated in this study: electron diffusion across magnetic fields and double ion production in inert-gas thrusters. A model is developed to describe electron diffusion across a magnetic field that is driven by both density and potential gradients, with Bohm diffusion used to predict the diffusion rate. This model has applications to conduction across magnetic fields inside a discharge chamber, as well as through a magnetic baffle region used to isolate a hollow cathode from the main chamber. A theory for double ion production is presented, which is not as complete as the electron diffusion theory described, but it should be a useful tool for predicting double ion sputter erosion. Correlations are developed that may be used, without experimental data, to predict double ion densities for the design of new and especially larger ion thrusters.

  14. Second Magnetoplasmadynamic Thruster Workshop

    NASA Technical Reports Server (NTRS)

    1992-01-01

    The meeting focused on progress made in establishing performance and lifetime expectations of magnetoplasmadynamic (MPD) thrusters as functions of power, propellant, and design; models for the plasma flow and electrode components; viability and transportability of quasi-steady thruster testing; engineering requirements for high power, long life thrusters; and facilities and their requirements for performance and life testing.

  15. Preliminary Results of Plasma Flow Measurements in a 2 KW Segmented Hall Thruster

    SciTech Connect

    Y. Raitses; D. Staack; A. Dunaevsky; L. Dorf; N.J. Fisch

    2003-03-01

    A 2-kW Hall thruster was developed, built, and operated in an upgraded vacuum facility. The thruster performance and parameters of the plasma flow were measured by new diagnostics for plume measurements and plasma measurements inside the thruster channel. The thruster demonstrated efficient operation in terms of propellant and current utilization efficiencies in the input power range of 0.5-3.5 kW. Preliminary measurements of the ion energy spectra from the thruster axis region and the distribution of plasma parameters in the vicinity of the thruster exit are reported.

  16. High Throughput 600 Watt Hall Effect Thruster for Space Exploration

    NASA Technical Reports Server (NTRS)

    Szabo, James; Pote, Bruce; Tedrake, Rachel; Paintal, Surjeet; Byrne, Lawrence; Hruby, Vlad; Kamhawi, Hani; Smith, Tim

    2016-01-01

    A nominal 600-Watt Hall Effect Thruster was developed to propel unmanned space vehicles. Both xenon and iodine compatible versions were demonstrated. With xenon, peak measured thruster efficiency is 46-48% at 600-W, with specific impulse from 1400 s to 1700 s. Evolution of the thruster channel due to ion erosion was predicted through numerical models and calibrated with experimental measurements. Estimated xenon throughput is greater than 100 kg. The thruster is well sized for satellite station keeping and orbit maneuvering, either by itself or within a cluster.

  17. A Preliminary Investigation of Hall Thruster Technology

    NASA Technical Reports Server (NTRS)

    Gallimore, Alec D.

    1997-01-01

    A three-year NASA/BMDO-sponsored experimental program to conduct performance and plume plasma property measurements on two Russian Stationary Plasma Thrusters (SPTs) has been completed. The program utilized experimental facilitates at the University of Michigan's Plasmadynamics and Electric Propulsion Laboratory (PEPL). The main features of the proposed effort were as follows: (1) Characterized Hall thruster (and arcjet) performance by measuring ion exhaust velocity with probes at various thruster conditions; (2) Used a variety of probe diagnostics in the thruster plume to measure plasma properties and flow properties including T(sub e) and n(sub e) ion current density and ion energy distribution, and electric fields by mapping plasma potential; (3) Used emission spectroscopy to identify species within the plume and to measure electron temperatures. A key and unique feature of our research was our collaboration with Russian Hall thruster researcher Dr. Sergey A Khartov, Deputy Dean of International Relations at the Moscow Aviation Institute (MAI). His activities in this program included consulting on and participation in research at PEPL through use of a MAI-built SPT and ion energy probe.

  18. Los Alamos NEP research in advanced plasma thrusters

    NASA Technical Reports Server (NTRS)

    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.

  19. The evolutionary development of high specific impulse electric thruster technology

    NASA Technical Reports Server (NTRS)

    Sovey, James S.; Hamley, John A.; Patterson, Michael J.; Rawlin, Vincent K.; Myers, Roger M.

    1992-01-01

    Electric propulsion flight and technology demonstrations conducted primarily by Europe, Japan, China, the U.S., and the USSR are reviewed. Evolutionary mission applications for high specific impulse electric thruster systems are discussed, and the status of arcjet, ion, and magnetoplasmadynamic thrusters and associated power processor technologies are summarized.

  20. Electric thruster models for multimegawatt nuclear electric propulsion mission design

    NASA Technical Reports Server (NTRS)

    Leifer, Stephanie D.; Blandino, John J.; Sercel, Joel C.

    1991-01-01

    Three types of electric thrusters currently under development at JPL have potential to support future missions which utilize multimegawatt nuclear electric propulsion. These electric thrusters are the electron bombardment ion thruster, the magnetoplasmadynamic (MPD) thruster, and the electron-cyclotron-resonance (ECR) thruster. The electron bombardment ion thruster is a relatively mature technology which has been developed for operation at kilowatt power levels but will require new development for application in the multimegawatt regime. The MPD engine represents a technology which may be very well suited to steady-state multimegawatt applications but which has been limited to sub-scale (100's of kW) and pulsed (MW) testing thus far. The ECR plasma engine represents a class of very promising new concepts which are still in the basic research phase of development, but which may possess important fundamental advantages over other electric thruster technologies. Models of these thrusters are described and used to make projections of thrusters specific mass, efficiency, and power handling capacity for operation in the multimegawatt regime.

  1. Experimental studies of an ECR plasma thruster

    NASA Technical Reports Server (NTRS)

    Kaufman, D. A.; Goodwin, D. G.; Sercel, J. C.

    1993-01-01

    The Electron Cyclotron Resonance (ECR) thruster is a proposed electrodeless space electric propulsion device with interesting and little understood physics. A laboratory ECR thruster was run in a vacuum tank at pressures in the 10 exp -5 torr range using 2.12 GHz microwave beam and Ar gas propellant. Movable diagnostic probes (a Faraday cup and a gridded energy analyzer) measured plasma characteristics as propellant gas flow rate and input microwave power level were varied. Ion energy and flux data were used to calculate I(sp), propulsive efficiency, and thrust. The ion flux profiles show an unexpected depression on the thruster axis for low tank pressures that disappears as the tank pressure increases. Ion energies decrease as the flow rate and pressure increase, but the microwave power level affects the energy only negligibly. The calculated propulsion parameters demonstrate that the efficiency of the laboratory device is low, and that tank pressure greatly changes the performance.

  2. Experimental studies of an ECR plasma thruster

    NASA Astrophysics Data System (ADS)

    Kaufman, D. A.; Goodwin, D. G.; Sercel, J. C.

    1993-06-01

    The Electron Cyclotron Resonance (ECR) thruster is a proposed electrodeless space electric propulsion device with interesting and little understood physics. A laboratory ECR thruster was run in a vacuum tank at pressures in the 10 exp -5 torr range using 2.12 GHz microwave beam and Ar gas propellant. Movable diagnostic probes (a Faraday cup and a gridded energy analyzer) measured plasma characteristics as propellant gas flow rate and input microwave power level were varied. Ion energy and flux data were used to calculate I(sp), propulsive efficiency, and thrust. The ion flux profiles show an unexpected depression on the thruster axis for low tank pressures that disappears as the tank pressure increases. Ion energies decrease as the flow rate and pressure increase, but the microwave power level affects the energy only negligibly. The calculated propulsion parameters demonstrate that the efficiency of the laboratory device is low, and that tank pressure greatly changes the performance.

  3. Advanced space propulsion thruster research

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1981-01-01

    Experiments showed that stray magnetic fields can adversely affect the capacity of a hollow cathode neutralizer to couple to an ion beam. Magnetic field strength at the neutralizer cathode orifice is a crucial factor influencing the coupling voltage. The effects of electrostatic accelerator grid aperture diameters on the ion current extraction capabilities were examined experimentally to describe the divergence, deflection, and current extraction capabilities of grids with the screen and accelerator apertures displaced relative to one another. Experiments performed in orificed, mercury hollow cathodes support the model of field enhanced thermionic electron mission from cathode inserts. Tests supported the validity of a thermal model of the cathode insert. A theoretical justification of a Saha equation model relating cathode plasma properties is presented. Experiments suggest that ion loss rates to discharge chamber walls can be controlled. A series of new discharge chamber magnetic field configurations were generated in the flexible magnetic field thruster and their effect on performance was examined. A technique used in the thruster to measure ion currents to discharge chamber walls is described. Using these ion currents the fraction of ions produced that are extracted from the discharge chamber and the energy cost of plasma ions are computed.

  4. VHITAL-160 Thruster Development Status

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Marrese-Reading, Colleen; Hofer, Rich; Owens, Al; Swindlehurst, Ray; Fitzgerald, Dennis

    2006-01-01

    A general overview on the status of the Very High Isp Thruster with Anode Layer (VHITAL)-160 program is presented. The topics include: 1) Bi TAL Overview; 2) VHITAL Program Overview; 3) Thruster Fabrication; and 4) Thruster Testing.

  5. Investigation of mercury thruster isolators

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1973-01-01

    Mercury ion thruster isolator lifetime tests were performed using different isolator materials and geometries. Tests were performed with and without the flow of mercury through the isolators in an oil diffusion pumped vacuum facility and cryogenically pumped bell jar. The onset of leakage current in isolators occurred in time intervals ranging from a few hours to many hundreds of hours. In all cases, surface contamination was responsible for the onset of leakage current and subsequent isolator failure. Rate of increase of leakage current and the leakage current level increased approximately exponentially with isolator temperature. Careful attention to shielding techniques and the elimination of sources of metal oxides appear to have eliminated isolator failures as a thruster life limiting mechanism.

  6. High Power Hall Thrusters

    NASA Technical Reports Server (NTRS)

    Jankovsky, Robert; Tverdokhlebov, Sergery; Manzella, David

    1999-01-01

    The development of Hall thrusters with powers ranging from tens of kilowatts to in excess of one hundred kilowatts is considered based on renewed interest in high power. high thrust electric propulsion applications. An approach to develop such thrusters based on previous experience is discussed. It is shown that the previous experimental data taken with thrusters of 10 kW input power and less can be used. Potential mass savings due to the design of high power Hall thrusters are discussed. Both xenon and alternate thruster propellant are considered, as are technological issues that will challenge the design of high power Hall thrusters. Finally, the implications of such a development effort with regard to ground testing and spacecraft intecrati'on issues are discussed.

  7. Pulsed hall thruster system

    NASA Technical Reports Server (NTRS)

    Hruby, Vladimir J. (Inventor); Pote, Bruce M. (Inventor); Gamero-Castano, Manuel (Inventor)

    2004-01-01

    A pulsed Hall thruster system includes a Hall thruster having an electron source, a magnetic circuit, and a discharge chamber; a power processing unit for firing the Hall thruster to generate a discharge; a propellant storage and delivery system for providing propellant to the discharge chamber and a control unit for defining a pulse duration .tau.<0.1d.sup.3.rho./m, where d is the characteristic size of the thruster, .rho. is the propellant density at standard conditions, and m is the propellant mass flow rate for operating either the power processing unit to provide to the Hall thruster a power pulse of a pre-selected duration, .tau., or operating the propellant storage and delivery system to provide a propellant flow pulse of duration, .tau., or providing both as pulses, synchronized to arrive coincidentally at the discharge chamber to enable the Hall thruster to produce a discreet output impulse.

  8. MPD thruster technology

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.

    1991-01-01

    Inhouse magnetoplasmadynamic (MPD) thruster technology is discussed. The study focussed on steady state thrusters at powers of less than 1 MW. Performance measurement and diagnostics technologies were developed for high power thrusters. Also developed was a MPD computer code. The stated goals of the program are to establish: performance and life limitation; influence of applied fields; propellant effects; and scaling laws. The presentation is mostly through graphs and charts.

  9. Conducting Wall Hall Thrusters

    NASA Technical Reports Server (NTRS)

    Goebel, Dan M.; Hofer, Richard R.; Mikellides, Ioannis G.; Katz, Ira; Polk, James E.; Dotson, Brandon

    2013-01-01

    A unique configuration of the magnetic field near the wall of Hall thrusters, called Magnetic Shielding, has recently demonstrated the ability to significantly reduce the erosion of the boron nitride (BN) walls and extend the life of Hall thrusters by orders of magnitude. The ability of magnetic shielding to minimize interactions between the plasma and the discharge chamber walls has for the first time enabled the replacement of insulating walls with conducting materials without loss in thruster performance. The boron nitride rings in the 6 kW H6 Hall thruster were replaced with graphite that self-biased to near the anode potential. The thruster efficiency remained over 60% (within two percent of the baseline BN configuration) with a small decrease in thrust and increase in Isp typical of magnetically shielded Hall thrusters. The graphite wall temperatures decreased significantly compared to both shielded and unshielded BN configurations, leading to the potential for higher power operation. Eliminating ceramic walls makes it simpler and less expensive to fabricate a thruster to survive launch loads, and the graphite discharge chamber radiates more efficiently which increases the power capability of the thruster compared to conventional Hall thruster designs.

  10. Oxygen-Methane Thruster

    NASA Technical Reports Server (NTRS)

    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.

  11. Diagnostics Systems for Permanent Hall Thrusters Development

    NASA Astrophysics Data System (ADS)

    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

  12. NASA's Hall Thruster Program

    NASA Technical Reports Server (NTRS)

    Jankovsky, Robert S.; Jacobson, David T.; Rawlin, Vincent K.; Mason, Lee S.; Mantenieks, Maris A.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

    2001-01-01

    NASA's Hall thruster program has base research and focused development efforts in support of the Advanced Space Transportation Program, Space-Based Program, and various other programs. The objective of the base research is to gain an improved understanding of the physical processes and engineering constraints of Hall thrusters to enable development of advanced Hall thruster designs. Specific technical questions that are current priorities of the base effort are: (1) How does thruster life vary with operating point? (2) How can thruster lifetime and wear rate be most efficiently evaluated? (3) What are the practical limitations for discharge voltage as it pertains to high specific impulse operation (high discharge voltage) and high thrust operation (low discharge voltage)? (4) What are the practical limits for extending Hall thrusters to very high input powers? and (5) What can be done during thruster design to reduce cost and integration concerns? The objective of the focused development effort is to develop a 50 kW-class Hall propulsion system, with a milestone of a 50 kW engineering model thruster/system by the end of program year 2006. Specific program wear 2001 efforts, along with the corporate and academic participation, are described.

  13. The flexible magnetic field thruster

    NASA Technical Reports Server (NTRS)

    Brophy, J. R.; Wilbur, P. J.

    1982-01-01

    The thruster is designed so that ion currents to various internal surfaces can be measured directly; these measurements facilitate calculations of the distribution of ion currents inside the discharge chamber. Experiments are described suggesting that the distribution of ion currents inside the discharge chamber is strongly dependent on the shape and strength of the magnetic field but independent of the discharge current, discharge voltage, and neutral flow rate. Measurements of the energy cost per plasma ion suggest that this cost decreases with increasing magnetic field strength as a consequence of increased anode shielding from the primary electrons. Energy costs per argon plasma ion as low as 50 eV are measured. The energy cost per beam ion is found to be a function of the energy cost per plasma ion, extracted ion fraction, and discharge voltage. Part of the energy cost per beam ion has to do with creating many ions in the plasma and then extracting only a fraction of them into the beam. The balance of the energy goes into accelerating the remaining plasma ions into the walls of the discharge chamber.

  14. Pulsed Inductive Thruster (PIT) Clamped Discharge Evaluation

    DTIC Science & Technology

    1988-12-31

    Lovberg, "Pulsed Inductive Thruster Technology", AFAL TR-87-012, April 1987. [2] Lyman Spitzer , Jr . "The Physics of Fully Ionized Gases" Interscience, New...Plasma (electron-ion) resistivity has the Spitzer dependence on temperature, with the coulomb factor In A equal to approximately 5: r7P = 10-3T-1 1

  15. Design and Preliminary Performance Testing of Electronegative Gas Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Liu, Thomas M.; Schloeder, Natalie R.; Walker, Mitchell L. R.; Polzin, Kurt A.; Dankanich, John W.; Aanesland, Ane

    2014-01-01

    In classical gridded electrostatic ion thrusters, positively charged ions are generated from a plasma discharge of noble gas propellant and accelerated to provide thrust. To maintain overall charge balance on the propulsion system, a separate electron source is required to neutralize the ion beam as it exits the thruster. However, if high-electronegativity propellant gases (e.g., sulfur hexafluoride) are instead used, a plasma discharge can result consisting of both positively and negatively charged ions. Extracting such electronegative plasma species for thrust generation (e.g., with time-varying, bipolar ion optics) would eliminate the need for a separate neutralizer cathode subsystem. In addition for thrusters utilizing a RF plasma discharge, further simplification of the ion thruster power system may be possible by also using the RF power supply to bias the ion optics. Recently, the PEGASES (Plasma propulsion with Electronegative gases) thruster prototype successfully demonstrated proof-of-concept operations in alternatively accelerating positively and negatively charged ions from a RF discharge of a mixture of argon and sulfur hexafluoride.i In collaboration with NASA Marshall Space Flight Center (MSFC), the Georgia Institute of Technology High-Power Electric Propulsion Laboratory (HPEPL) is applying the lessons learned from PEGASES design and testing to develop a new thruster prototype. This prototype will incorporate design improvements and undergo gridless operational testing and diagnostics checkout at HPEPL in April 2014. Performance mapping with ion optics will be conducted at NASA MSFC starting in May 2014. The proposed paper discusses the design and preliminary performance testing of this electronegative gas plasma thruster prototype.

  16. Colloid Thrusters, Physics, Fabrication and Performance

    DTIC Science & Technology

    2005-11-17

    Murray, Paulo Lozano and Manuel Martinez-Sanchez, "Mass Spectrometric Analysis of Colloid Thruster Ion Emission from Selected Propellants", JOURNAL...OF PROPULSION AND POWER Vol. 21, No. 3, May-June 2005. 21. Paulo Lozano and Manuel Martinez-Sanchez, "Ionic liquid ion sources: characterization of...externally wetted emitters", Journal of Colloid and Interface Science 282 (2005) 415-421. 22. Paulo Lozano and Manuel Martinez-Sanchez, "Efficiency

  17. Miniature cold gas thrusters

    NASA Astrophysics Data System (ADS)

    Bzibziak, R. J., Sr.

    1992-07-01

    Cold gas thrusters provide a safe, inexpensive, lightweight and reliable means of propulsive control for small satellites, projectiles and maneuvering control systems. Moog Inc. has designed and developed a family of miniature cold gas thrusters for use on Strategic Defense Iniative flight simulation experiments, sounding rockets, small satellite applications, astronaut control systems, and close proximity maneuvering systems for Space System. Construction features such as coil assembly, core assembly, armature assembly, external housing and valve body are discussed. The design approach, performance characteristics and functional description of cold gas thrusters designed for various applications are presented.

  18. Global model of an iodine gridded plasma thruster

    NASA Astrophysics Data System (ADS)

    Grondein, P.; Lafleur, T.; Chabert, P.; Aanesland, A.

    2016-03-01

    Most state-of-the-art electric space propulsion systems such as gridded and Hall effect thrusters use xenon as the propellant gas. However, xenon is very rare, expensive to produce, and used in a number of competing industrial applications. Alternatives to xenon are currently being investigated, and iodine has emerged as a potential candidate. Its lower cost and larger availability, its solid state at standard temperature and pressure, its low vapour pressure and its low ionization potential make it an attractive option. In this work, we compare the performances of a gridded ion thruster operating separately with iodine and xenon, under otherwise identical conditions using a global model. The thruster discharge properties such as neutral, ion, and electron densities and electron temperature are calculated, as well as the thruster performance parameters such as thrust, specific impulse, and system efficiencies. For similar operating conditions, representative of realistic thrusters, the model predicts similar thrust levels and performances for both iodine and xenon. The thruster efficiency is however slightly higher for iodine compared with xenon, due to its lower ionization potential. This demonstrates that iodine could be a viable alternative propellant for gridded plasma thrusters.

  19. Integrated thruster assembly program

    NASA Technical Reports Server (NTRS)

    1973-01-01

    The program is reported which has provided technology for a long life, high performing, integrated ACPS thruster assembly suitable for use in 100 typical flights of a space shuttle vehicle over a ten year period. The four integrated thruster assemblies (ITA) fabricated consisted of: propellant injector; a capacitive discharge, air gap torch type igniter assembly; fast response igniter and main propellant valves; and a combined regen-dump film cooled chamber. These flightweight 6672 N (1500 lb) thruster assemblies employed GH2/GO2 as propellants at a chamber pressure of 207 N/sq cm (300 psia). Test data were obtained on thrusted performance, thermal and hydraulic characteristics, dynamic response in pulsing, and cycle life. One thruster was fired in excess of 42,000 times.

  20. MPD thruster technology

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Mantenieks, Maris A.; Lapointe, Michael R.

    1991-01-01

    MPD (MagnetoPlasmaDynamic) thrusters demonstrated between 2000 and 7000 seconds specific impulse at efficiencies approaching 40 percent, and were operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. To date, however, only a limited number of thruster configurations, propellants, and operating conditions were studied. The present status of MPD research is reviewed, including developments in the measured performance levels and electrode erosion rates. Theoretical studies of the thruster dynamics are also described. Significant progress was made in establishing empirical scaling laws, performance and lifetime limitations and in the development of numerical codes to simulate the flow field and electrode processes.

  1. MPD thruster technology

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Lapointe, Michael R.; Mantenieks, Maris A.

    1991-01-01

    MPD thrusters have demonstrated between 2000 and 7000 sec specific impulse at efficiencies approaching 40 percent, and have been operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. This work reviews the present status of MPD thruster research, including developments in the measured performance levels and electrode erosion rates, and theoretical studies of the thruster dynamics. Significant progress has been made in establishing empirical scaling laws, performance and lifetime limitations, and in the development of numerical codes to simulate the flowfield and the electrode processes.

  2. Development of an 8-cm engineering model thruster system

    NASA Technical Reports Server (NTRS)

    Herron, B. G.; Hyman, J., Jr.; Hopper, D. J.

    1976-01-01

    Electric propulsion has been shown to offer major advantages over the techniques currently employed for the control of earth satellites. For a user to realize these advantages, however, requires the availability of a proven, operationally flight-ready propulsion system. Currently an Engineering Model of an 8-cm ion thruster propulsion system is under development. The system includes the thruster unit with its associated reservoir, thruster gimbaling subsystem, and power processing unit. This paper describes the EM System with special emphasis on hardware design and system performance.

  3. Controlling the Plasma Flow in the Miniaturized Cylindrical Hall Thruster

    SciTech Connect

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

    2008-03-04

    A substantial narrowmg of the plume of the cylindrical RaIl thruster (CRT) was observed upon the enhancement of the electron emission from the hollow cathode discharge, which implies the possibility for the thruster efficiency increase due to the ion beam focusing. It is demonstrated that the miniaturized CRT can be operated in the non-self-sustained regime, with the discharge current limited by the cathode electron emission. The thruster operation in this mode greatly expands the range of the plasma and discharge parameters normally accessible for the CRT.

  4. Acceleration Mechanism Of Pulsed Laser-Electromagnetic Hybrid Thruster

    SciTech Connect

    Horisawa, Hideyuki; Mashima, Yuki; Yamada, Osamu

    2011-11-10

    A fundamental study of a newly developed rectangular pulsed laser-electromagnetic hybrid thruster was conducted. Laser-ablation plasma in the thruster was induced through laser beam irradiation onto a solid target and accelerated by electrical means instead of direct acceleration only by using a laser beam. The performance of the thrusters was evaluated by measuring the ablated mass per pulse and impulse bit. As results, significantly high specific impulses up to 7,200 s were obtained at charge energies of 8.6 J. Moreover, from the Faraday cup measurement, it was confirmed that the speed of ions was accelerated with addition of electric energy.

  5. Metallic Wall Hall Thrusters

    NASA Technical Reports Server (NTRS)

    Goebel, Dan Michael (Inventor); Hofer, Richard Robert (Inventor); Mikellides, Ioannis G. (Inventor)

    2016-01-01

    A Hall thruster apparatus having walls constructed from a conductive material, such as graphite, and having magnetic shielding of the walls from the ionized plasma has been demonstrated to operate with nearly the same efficiency as a conventional non-magnetically shielded design using insulators as wall components. The new design is believed to provide the potential of higher power and uniform operation over the operating life of a thruster device.

  6. Characterization of an electrodeless ECR plasma thruster

    NASA Astrophysics Data System (ADS)

    Vialis, Théo; Jarrige, Julien; Packan, Denis

    2016-09-01

    Several advanced plasma thruster technologies are currently being studied for the 1-10 mN range. ONERA is developing an Electron Cyclotron Resonance (ECR) plasma thruster, whose main advantage is to produce a current-free plume. It does not need a neutralizing cathode, which is one of the most fragile component in electrostatic thrusters. The ECR thruster consists of a coaxial structure immersed in an axial divergent magnetic field, fed with xenon. A plasma is generated by resonant absorption of microwave power (at 2.45 GHz) and is accelerated in an electron driven magnetic nozzle to produce the thrust. Previous measurements, performed with electrostatic probes, have shown promising performances. Electrons are heated at very high temperatures (several tens of eV), and ion kinetic energy is up to 400 eV in the plume. The estimated thrust is 1 mN, with an efficiency of 16%, for a power of 30W. In this work, a new version of the device has been conceived for direct thrust measurement on a dedicated thrust balance. The effect of magnetic field topology, propellant, mass flow rate and absorbed power are investigated. Thrust measurement are compared with values estimated from electrostatic probes results (ion current and energy).

  7. 43. Bow thruster room. Bow thruster engine not used for ...

    Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

    43. Bow thruster room. Bow thruster engine not used for powering hydraulics to boom as in some other tenders in same class. - U.S. Coast Guard Cutter BRAMBLE, Waterfront at Lincoln Avenue, Port Huron, St. Clair County, MI

  8. Mode Transitions in Hall Effect Thrusters

    NASA Technical Reports Server (NTRS)

    Sekerak, Michael J.; Longmier, Benjamin W.; Gallimore, Alec D.; Brown, Daniel L.; Hofer, Richard R.; Polk, James E.

    2013-01-01

    Mode transitions have been commonly observed in Hall Effect Thruster (HET) operation where a small change in a thruster operating parameter such as discharge voltage, magnetic field or mass flow rate causes the thruster discharge current mean value and oscillation amplitude to increase significantly. Mode transitions in a 6-kW-class HET called the H6 are induced by varying the magnetic field intensity while holding all other operating parameters constant and measurements are acquired with ion saturation probes and ultra-fast imaging. Global and local oscillation modes are identified. In the global mode, the entire discharge channel oscillates in unison and azimuthal perturbations (spokes) are either absent or negligible. Downstream azimuthally spaced probes show no signal delay between each other and are very well correlated to the discharge current signal. In the local mode, signals from the azimuthally spaced probes exhibit a clear delay indicating the passage of "spokes" and are not well correlated to the discharge current. These spokes are localized oscillations propagating in the ExB direction that are typically 10-20% of the mean value. In contrast, the oscillations in the global mode can be 100% of the mean value. The transition between global and local modes occurs at higher relative magnetic field strengths for higher mass flow rates or higher discharge voltages. The thrust is constant through mode transition but the thrust-to-power decreased by 25% due to increasing discharge current. The plume shows significant differences between modes with the global mode significantly brighter in the channel and the near-field plasma plume as well as exhibiting a luminous spike on thruster centerline. Mode transitions provide valuable insight to thruster operation and suggest improved methods for thruster performance characterization.

  9. NASA's Hall Thruster Program 2002

    NASA Technical Reports Server (NTRS)

    Jankovsky, Robert S.; Jacobson, David T.; Pinero, Luis R.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

    2002-01-01

    The NASA Hall thruster program currently supports a number of tasks related to high power thruster development for a number of customers including the Energetics Program (formerly called the Space-based Program), the Space Solar Power Program, and the In-space Propulsion Program. In program year 2002, two tasks were central to the NASA Hall thruster program: 1) the development of a laboratory Hall thruster capable of providing high thrust at high power-, and 2) investigations into operation of Hall thrusters at high specific impulse. In addition to these two primary thruster development activities, there are a number of other on-going activities supported by the NASA Hall thruster program. These additional activities are related to issues such as high-power power processor architecture, thruster lifetime, and spacecraft integration.

  10. Change in transmittance of fused silica as a means of detecting material sputtered from components on a 5-cm ion thruster

    NASA Technical Reports Server (NTRS)

    Weigand, A. J.; Mirtich, M. J.

    1972-01-01

    Two endurance tests of a 5-cm mercury bombardment thruster are reported. Both tests used a translational screen-grid system with the beam vectored 10 degrees. The first test lasted 141 hours and the second test operated for 2026 hours. In each test two fused silica samples (solar cell covers), 2.0 cm by 2.1 cm, were placed in shielded holders to detect materials sputtered from the thruster. Spectral optical properties between 0.398 and 2.16 microns were measured on each sample, both before and after the endurance tests. The deposition on each sample was spectrographically analyzed to determine the type of materials sputtered from the thruster. It was found that sputtering from the neutralizer is highly dependent on its position with respect to the beam edge. The sputtering from the accelerator grid of the translational screen-grid system of the 2026 hour test was sufficient to form an opaque film on the sample located in the direction opposite to the vectored beam.

  11. Iodine Hall Thruster

    NASA Technical Reports Server (NTRS)

    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).

  12. Development Status of the NSTAR Ion Propulsion System Power Processor

    NASA Technical Reports Server (NTRS)

    Hamley, John A.; Pinero, Luis R.; Rawlin, Vincent K.; Miller, John R.; Cartier, Kevin C.; Bowers, Glen E.

    1995-01-01

    A 0.5-2.3 kW xenon ion propulsion system is presently being developed under the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program. This propulsion system includes a 30 cm diameter xenon ion thruster, a Digital Control Interface Unit, a xenon feed system, and a power processing unit (PPU). The PPU consists of the power supply assemblies which operate the thruster neutralizer, main discharge chamber, and ion optics. Also included are recycle logic and a digital microcontroller. The neutralizer and discharge power supplies employ a dual use configuration which combines the functions of two power supplies into one, significantly simplifying the PPU. Further simplification was realized by implementing a single thruster control loop which regulates the beam current via the discharge current. Continuous throttling is possible over a 0.5-2.3 kW output power range. All three power supplies have been fabricated and tested with resistive loads, and have been combined into a single breadboard unit with the recycle logic and microcontroller. All line and load regulation test results show the power supplies to be within the NSTAR flight PPU specified power output of 1.98 kW. The overall efficiency of the PPU, calculated as the combined efficiencies of the power supplies and controller, at 2.3 kW delivered to resistive loads was 0.90. The component was 6.16 kg. Integration testing of the neutralizer and discharge power supplies with a functional model thruster revealed no issues with discharge ignition or steady state operation.

  13. A collisionless plasma thruster plume expansion model

    NASA Astrophysics Data System (ADS)

    Merino, Mario; Cichocki, Filippo; Ahedo, Eduardo

    2015-06-01

    A two-fluid model of the unmagnetized, collisionless far region expansion of the plasma plume for gridded ion thrusters and Hall effect thrusters is presented. The model is integrated into two semi-analytical solutions valid in the hypersonic case. These solutions are discussed and compared against the results from the (exact) method of characteristics; the relative errors in density and velocity increase slowly axially and radially and are of the order of 10-2-10-3 in the cases studied. The plasma density, ion flux and ambipolar electric field are investigated. A sensitivity analysis of the problem parameters and initial conditions is carried out in order to characterize the far plume divergence angle in the range of interest for space electric propulsion. A qualitative discussion of the physics of the secondary plasma plume is also provided.

  14. NASA's 2004 Hall Thruster Program

    NASA Technical Reports Server (NTRS)

    Jacobson, David T.; Manzella, David H.; Hofer, Richard R.; Peterson, Peter Y.

    2004-01-01

    An overview of NASA's Hall thruster research and development tasks conducted during fiscal year 2004 is presented. These tasks focus on: raising the technology readiness level of high power Hall thrusters, developing a moderate-power/ moderate specific impulse Hall thruster, demonstrating high-power/high specific impulse Hall thruster operation, and addressing the fundamental technical challenges of emerging Hall thruster concepts. Programmatic background information, technical accomplishments and out year plans for each program element performed under the sponsorship of the In-Space Transportation Program, Project Prometheus, and the Energetics Project are provided.

  15. Optimization of a wall-less Hall thruster

    NASA Astrophysics Data System (ADS)

    Vaudolon, Julien; Mazouffre, Stéphane; Hénaux, Carole; Harribey, Dominique; Rossi, Alberto

    2015-10-01

    An experimental optimization of a Hall thruster in wall-less operation mode is performed with the PPS-Flex, a 1.5 kW class thruster capable of modifying the magnetic field topology over a broad range of configurations. The anode geometry and the magnetic topology have been modified to avoid interaction between the magnetic field lines and the anode surface, compared to the first wall-less Hall thruster prototype. The measurements of the thrust and far-field ion properties reveal that a satisfactory performance level can be obtained once the magnetic barrier is restored, and pave the way towards the development of a high-efficiency wall-less Hall thruster.

  16. Performance of a Low-Power Cylindrical Hall Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Markusic, Thomas E.; Stanojev, Boris J.; Dehoyos, Amado; Raitses, Yevgeny; Smirnov, Artem; Fisch, Nathaniel J.

    2007-01-01

    Recent mission studies have shown that a Hall thruster which operates at relatively constant thrust efficiency (45-55%) over a broad power range (300W - 3kW) is enabling for deep space science missions when compared with slate-of-the-art ion thrusters. While conventional (annular) Hall thrusters can operate at high thrust efficiency at kW power levels, it is difficult to construct one that operates over a broad power envelope down to 0 (100 W) while maintaining relatively high efficiency. In this note we report the measured performance (I(sub sp), thrust and efficiency) of a cylindrical Hall thruster operating at 0 (100 W) input power.

  17. Magnesium Hall Thruster

    NASA Technical Reports Server (NTRS)

    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.

  18. Magnetoplasmadynamic Thruster Workshop

    NASA Technical Reports Server (NTRS)

    1991-01-01

    On May 16, 1991, the NASA Headquarters Propulsion, Power, and Energy Division and the NASA Lewis Research Center Low Thrust Propulsion Branch hosted a workshop attended by key experts in magnetoplasmadynamic (MPD) thrusters and associated sciences. The scope was limited to high power MPD thrusters suitable for major NASA space exploration missions, and its purpose was to initiate the process of increasing the expectations and prospects for MPD research, primarily by increasing the level of cooperation, interaction, and communication between parties within the MPD community.

  19. Low Cost Electric Propulsion Thruster for Deep Space Robotic Science Missions

    NASA Technical Reports Server (NTRS)

    Manzella, David

    2008-01-01

    Electric Propulsion (EP) has found widespread acceptance by commercial satellite providers for on-orbit station keeping due to the total life cycle cost advantages these systems offer. NASA has also sought to benefit from the use of EP for primary propulsion onboard the Deep Space-1 and DAWN spacecraft. These applications utilized EP systems based on gridded ion thrusters, which offer performance unequaled by other electric propulsion thrusters. Through the In-Space Propulsion Project, a lower cost thruster technology is currently under development designed to make electric propulsion intended for primary propulsion applications cost competitive with chemical propulsion systems. The basis for this new technology is a very reliable electric propulsion thruster called the Hall thruster. Hall thrusters, which have been flown by the Russians dating back to the 1970s, have been used by the Europeans on the SMART-1 lunar orbiter and currently employed by 15 other geostationary spacecraft. Since the inception of the Hall thruster, over 100 of these devices have been used with no known failures. This paper describes the latest accomplishments of a development task that seeks to improve Hall thruster technology by increasing its specific impulse, throttle-ability, and lifetime to make this type of electric propulsion thruster applicable to NASA deep space science missions. In addition to discussing recent progress on this task, this paper describes the performance and cost benefits projected to result from the use of advanced Hall thrusters for deep space science missions.

  20. Primary electric propulsion technology study. [for thruster wear-out mechanisms

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.; Beattie, J. R.

    1979-01-01

    An investigation of the 30-cm engineering-model-thruster technology with emphasis placed on the development of models for understanding and predicting the operational characteristics and wear-out mechanisms of the thruster as a function of operating or design parameters is presented. The task studies include: (1) the wear mechanisms and wear rates that determine the useful lifetime of the thruster discharge chamber; (2) cathode lifetime as determined by the depletion of barium from the barium-aluminate-impregnated-porous-tungsten insert that serves as a barium reservoir; (3) accelerator-grid-system technology; (4) a verification of the high-voltage propellant-flow-electrical-isolator design developed under NASA contract NAS3-20395 for operation at 10-kV applied voltage and 10-A equivalent propellant flow with mercury and argon propellants. A model was formulated for predicting performance.

  1. MPD Thruster Performance Analytic Models

    NASA Technical Reports Server (NTRS)

    Gilland, James; Johnston, Geoffrey

    2007-01-01

    Magnetoplasmadynamic (MPD) thrusters are capable of accelerating quasi-neutral plasmas to high exhaust velocities using Megawatts (MW) of electric power. These characteristics make such devices worthy of consideration for demanding, far-term missions such as the human exploration of Mars or beyond. Assessment of MPD thrusters at the system and mission level is often difficult due to their status as ongoing experimental research topics rather than developed thrusters. However, in order to assess MPD thrusters utility in later missions, some adequate characterization of performance, or more exactly, projected performance, and system level definition are required for use in analyses. The most recent physical models of self-field MPD thrusters have been examined, assessed, and reconfigured for use by systems and mission analysts. The physical models allow for rational projections of thruster performance based on physical parameters that can be measured in the laboratory. The models and their implications for the design of future MPD thrusters are presented.

  2. MPD Thruster Performance Analytic Models

    NASA Technical Reports Server (NTRS)

    Gilland, James; Johnston, Geoffrey

    2003-01-01

    Magnetoplasmadynamic (MPD) thrusters are capable of accelerating quasi-neutral plasmas to high exhaust velocities using Megawatts (MW) of electric power. These characteristics make such devices worthy of consideration for demanding, far-term missions such as the human exploration of Mars or beyond. Assessment of MPD thrusters at the system and mission level is often difficult due to their status as ongoing experimental research topics rather than developed thrusters. However, in order to assess MPD thrusters utility in later missions, some adequate characterization of performance, or more exactly, projected performance, and system level definition are required for use in analyses. The most recent physical models of self-field MPD thrusters have been examined, assessed, and reconfigured for use by systems and mission analysts. The physical models allow for rational projections of thruster performance based on physical parameters that can be measured in the laboratory. The models and their implications for the design of future MPD thrusters are presented.

  3. Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase I: Numerical Simulations

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.

    2011-01-01

    In a proof-of-principle effort to demonstrate the feasibility of magnetically shielded (MS) Hall thrusters, an existing laboratory thruster has been modified with the guidance of physics-based numerical simulation. When operated at a discharge power of 6-kilowatts the modified thruster has been designed to reduce the total energy and flux of ions to the channel insulators by greater than 1 and greater than 3 orders of magnitude, respectively. The erosion rates in this MS thruster configuration are predicted to be at least 2-4 orders of magnitude lower than those in the baseline (BL) configuration. At such rates no detectable erosion is expected to occur.

  4. Experimental results of an iodine plasma in PEGASES gridded thruster

    NASA Astrophysics Data System (ADS)

    Grondein, Pascaline; Aanesland, Ane

    2015-09-01

    In the electric gridded thruster PEGASES, both positive and negative ions are expelled after extraction from an ion-ion plasma. This ion-ion plasma is formed downstream a localized magnetic field placed a few centimeters from the ionization region, trapping and cooling down the electron to allow a better attachment to an electronegative gas. For this thruster concept, iodine has emerged as the most attractive option. Heavy, under diatomic form and therefore good for high thrust, its low ionization threshold and high electronegativity lead to high ion-ion densities and low RF power. After the proof-of-concept of PEGASES using SF6 as propellant, we present here experimental results of an iodine plasma studied inside PEGASES thruster. At solid state at standard temperature and pressure, iodine is heated to sublimate, then injected inside the chamber where the neutral gas is heated and ionized. The whole injection system is heated to avoid deposition on surfaces and a mass flow controller allows a fine control on the neutral gas mass flow. A 3D translation stage inside the vacuum chamber allows volumetric plasma studies using electrostatic probes. The results are also compared with the global model dedicated to iodine as propellant for electric gridded thrusters. This work has been done within the LABEX Plas@par project, and received financial state aid managed by the Agence Nationale de la Recherche, as part of the programme ``Investissements d'avenir.''

  5. Status of the NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test After 30,352 Hours of Operation

    NASA Technical Reports Server (NTRS)

    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 30,352 hr of operation and processed 490 kg of xenon throughput--surpassing the NSTAR Extended Life Test hours demonstrated and 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.

  6. A bibliography of electrothermal thruster technology, 1984

    NASA Technical Reports Server (NTRS)

    Sovey, J. S.; Hardy, T. L.; Englehart, M.

    1986-01-01

    Electrothermal propulsion concepts are briefly discussed as an introduction to a bibliography and author index. Nearly 700 citations are given for resistojets, thermal arcjets, pulsed electrothermal thrusters, microwave heated devices, solar thermal thrusters, and laser thermal thrusters.

  7. Electronegative Gas Thruster - Direct Thrust Measurement Project

    NASA Technical Reports Server (NTRS)

    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.

  8. NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

    NASA Technical Reports Server (NTRS)

    Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

    2012-01-01

    The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar-electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced mission capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to qualify the thruster propellant throughput capability. The thruster has set electric propulsion records for the longest operating duration, highest propellant throughput, and most total impulse demonstrated. At the time of this publication, the NEXT LDT has surpassed 42,100 h of operation, processed more than 736 kg of xenon propellant, and demonstrated greater than 28.1 MN s total impulse. Thruster performance has been steady with negligible degradation. The NEXT thruster design has mitigated several lifetime limiting mechanisms encountered in the NSTAR design, including the NSTAR first failure mode, thereby drastically improving thruster capabilities. Component erosion rates and the progression of the predicted life-limiting erosion mechanism for the thruster compare favorably to pretest predictions based upon semi-empirical ion thruster models used in the thruster service life assessment. Service life model validation has been accomplished by the NEXT LDT. Assuming full-power operation until test article failure, the models and extrapolated erosion data predict penetration of the accelerator grid grooves after more than 45,000 hours of operation while processing over 800 kg of xenon propellant. Thruster failure due to degradation of the accelerator grid structural integrity is expected after

  9. NASA's Evolutionary Xenon Thruster (NEXT) Long-Duration Test as of 736 kg of Propellant Throughput

    NASA Technical Reports Server (NTRS)

    Shastry, Rohit; Herman, Daniel A.; Soulas, George C.; Patterson, Michael J.

    2012-01-01

    The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation solar-electric ion propulsion system with significant enhancements beyond the state-of-the-art NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) ion propulsion system to provide future NASA science missions with enhanced mission capabilities. A Long-Duration Test (LDT) was initiated in June 2005 to validate the thruster service life modeling and to qualify the thruster propellant throughput capability. The thruster has set electric propulsion records for the longest operating duration, highest propellant throughput, and most total impulse demonstrated. At the time of this publication, the NEXT LDT has surpassed 42,100 h of operation, processed more than 736 kg of xenon propellant, and demonstrated greater than 28.1 MN s total impulse. Thruster performance has been steady with negligible degradation. The NEXT thruster design has mitigated several lifetime limiting mechanisms encountered in the NSTAR design, including the NSTAR first failure mode, thereby drastically improving thruster capabilities. Component erosion rates and the progression of the predicted life-limiting erosion mechanism for the thruster compare favorably to pretest predictions based upon semi-empirical ion thruster models used in the thruster service life assessment. Service life model validation has been accomplished by the NEXT LDT. Assuming full-power operation until test article failure, the models and extrapolated erosion data predict penetration of the accelerator grid grooves after more than 45,000 hours of operation while processing over 800 kg of xenon propellant. Thruster failure due to degradation of the accelerator grid structural integrity is expected after groove penetration.

  10. Investigation of mercury thruster isolators. [service life

    NASA Technical Reports Server (NTRS)

    Mantenieks, M. A.

    1973-01-01

    Mercury ion thruster isolator lifetime tests were performed using different isolator materials and geometries. Tests were performed with and without the flow of mercury through the isolators in an oil diffusion pumped vacuum facility and cryogenically pumped bell jar. The onset of leakage current in isolators tested occurred in time intervals ranging from a few hours to many hundreds of hours. In all cases, surface contamination was responsible for the onset of leakage current and subsequent isolator failure. Rate of increase of leakage current and the leakage current level increased approximately exponentially with isolator temperature. Careful attention to shielding techniques and the elimination of sources of metal oxides appear to have eliminated isolator failures as a thruster life limiting mechanism.

  11. Thermal Modeling for Pulsed Inductive FRC Plasmoid Thrusters

    NASA Astrophysics Data System (ADS)

    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.

  12. The Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

    Eskridge, R.; Martin, Adam; Lee, Michael; Smith, James; Koelfgen, Syri

    2003-01-01

    This viewgraph presentation describes the overall Plasma Thruster Experiment (PTX), it's purpose and design, compact toroid propulsion, advantages and requirements of a plasmoid thruster, the projected efficiency, theta-pinch formation, a simulation of the PTX Coil/Bank Circuit using SPICE, the test firing of the PTX Capacitor Bank, PTX diagnostics, the excluded flux array, thruster simulations using MOQUI, and future work on the PTX.

  13. Proposal for Testing and Validation of Vacuum Ultra-Violet Atomic Laser-Induced Fluorescence as a Method to Analyze Carbon Grid Erosion in Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Stevens, Richard

    2003-01-01

    Previous investigation under award NAG3-25 10 sought to determine the best method of LIF to determine the carbon density in a thruster plume. Initial reports from other groups were ambiguous as to the number of carbon clusters that might be present in the plume of a thruster. Carbon clusters would certainly affect the ability to LIF; if they were the dominant species, then perhaps the LIF method should target clusters. The results of quadrupole mass spectroscopy on sputtered carbon determined that minimal numbers of clusters were sputtered from graphite under impact from keV Krypton. There were some investigations in the keV range by other groups that hinted at clusters, but at the time the proposal was presented to NASA, there was no data from low-energy sputtering available. Thus, the proposal sought to develop a method to characterize the population only of atoms sputtered from a graphite target in a test cell. Most of the ground work had been established by the previous two years of investigation. The proposal covering 2003 sought to develop an anti-Stokes Raman shifting cell to generate VUW light and test this cell on two different laser systems, ArF and YAG- pumped dye. The second goal was to measure the lowest detectable amounts of carbon atoms by 156.1 nm and 165.7 nm LIF. If equipment was functioning properly, it was expected that these goals would be met easily during the timeframe of the proposal, and that is the reason only modest funding was requested. The PI was only funded at half- time by Glenn during the summer months. All other work time was paid for by Whitworth College. The college also funded a student, Charles Shawley, who worked on the project during the spring.

  14. RF micro-discharge thruster

    NASA Astrophysics Data System (ADS)

    Dunaevsky, Alexander; Fisch, Nathaniel

    2004-11-01

    Propulsion devices for spacecrafts with masses of several tens to one hundred kilograms are in an increasing demand. These devices should provide thrust of a few mN and specific impulse of about 1000 s at the total power consumption of several tens of W. In search of an alternative solution for lower power range, we investigated an rf discharge initiated in a sub-millimeter capillary fed by a gaseous propellant. In such a discharge, it is possible to heat plasma electrons up to temperatures of ˜ 20-30 eV. Steep density drop at the open end of the capillary should be a reason of the formation of a double layer, were the discharge ions are accelerated to energies of ˜5Te. A laboratory prototype demonstrated stable operation at the argon flow rate of 4-10 sccm. The discharge was powered by a 2 MHz rf generator. Power consumption of the discharge was about 16 W. Ionization rate was moderate due to nonoptimal electrode configuration, which resulted in the propellant utilization of 6-11%. Relatively wide plume angle of ˜130 degrees indicates that the acceleration region is placed outside the capillary and has a convex shape. Stability and parameters of the discharge depends on the material of the capillary channel. Among advantages of the rf micro-discharge thruster are simplicity, small size, and absence of cathode-neutralizer. Being optimized, the rf micro-discharge thruster seems very promising propulsion device for sub-mN thrust range.

  15. Thruster sealing system and apparatus

    NASA Technical Reports Server (NTRS)

    Svejkovsky, Paul A. (Inventor)

    1992-01-01

    A thruster nozzle sealing system and apparatus is provided for protection of spacecraft thruster motors. The system includes a sealing plug, a sealing plug insertion tool, an outer cover, an outer cover attachment, and a ferry flight attachment. The sealing plug prevents moisture from entering the thruster engine so as to prevent valve failure. The attachments are interchangeably connectable with the sealing plug. The ferry flight attachment is used during air transportation of the spacecraft, and the outer cover attachment is used during storage and service of the spacecraft. The outer cover provides protection to the thruster nozzle from mechanical damage.

  16. Multimegawatt MPD thruster design considerations

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Parkes, James E.; Mantenieks, Maris A.

    1992-01-01

    Performance and lifetime requirements for multimegawatt magnetoplasmadynamic (MPD) thrusters were used to establish a baseline 2.5 MW thruster design. The chamber surface power deposition resulting from current conduction, plasma and surface radiation, and conduction from the hot plasma was then evaluated to establish the feasibility of thruster operation. It was determined that state of the art lithium heat pipes were adequate to cool the anode electrode, and that the liquid hydrogen propellant could be used to cool the applied field magnet, cathode, and backplate. Unresolved issues having an impact of thruster design are discussed to help focus future research.

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

    NASA Technical Reports Server (NTRS)

    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.

  18. Green Liquid Monopropellant Thruster

    NASA Technical Reports Server (NTRS)

    Joshi, Prakash B.

    2015-01-01

    Physical Sciences, Inc. (PSI), and Orbital Technologies Corporation (ORBITEC) are developing a unique chemical propulsion system for next-generation NASA science spacecraft and missions. The system is compact, lightweight, and can operate with high reliability over extended periods of time and under a wide range of thermal environments. The system uses a new storable, low-toxicity liquid monopropellant as its working fluid. In Phase I, the team demonstrated experimentally the critical ignition and combustion processes for the propellant and used the data to develop thruster design concepts. In Phase II, the team developed and demonstrated in the laboratory a proof-of-concept prototype thruster. A Phase III project is envisioned to develop a full-scale protoflight propulsion system applicable to a class of NASA missions.

  19. Comparison of Hall Thruster Plume Expansion Model with Experimental Data

    DTIC Science & Technology

    2006-05-23

    focus of this study, is a hybrid particle- in-cell ( PIC ) model that tracks particles along an unstructured tetrahedral mesh. * Research Engineer...measurements of the ion current density profile, ion energy distributions, and ion species fraction distributions using a nude Faraday probe, retarding...Vol.37 No.1. 6 Oh, D. and Hastings, D., “Three Dimensional PIC -DSMC Simulations of Hall Thruster Plumes and Analysis for Realistic Spacecraft

  20. Performance Characterization of a Three-Axis Hall Effect Thruster

    DTIC Science & Technology

    2010-12-01

    especially solar arrays. The ob- vious solution is to orient the thruster plume away from sensitive areas and to avoid the spacecraft entirely if...ion current at a point in the plasma. When ions strike the probe electrons move to the collector plate which is measured by the probe as the ion...and strike the collector plate. The current value is determined by measuring the voltage drop across a resistor which separates the collector plate

  1. Engineering model 8-cm thruster subsystem

    NASA Technical Reports Server (NTRS)

    Herron, B. G.; Hyman, J.; Hopper, D. J.; Williamson, W. S.; Dulgeroff, C. R.; Collett, C. R.

    1978-01-01

    An Engineering Model (EM) 8 cm Ion Thruster Propulsion Subsystem was developed for operation at a thrust level 5 mN (1.1 mlb) at a specific impulse 1 sub sp = 2667 sec with a total system input power P sub in = 165 W. The system dry mass is 15 kg with a mercury-propellant-reservoir capacity of 8.75 kg permitting uninterrupted operation for about 12,500 hr. The subsystem can be started from a dormant condition in a time less than or equal to 15 min. The thruster has a design lifetime of 20,000 hr with 10,000 startup cycles. A gimbal unit is included to provide a thrust vector deflection capability of + or - 10 degrees in any direction from the zero position. The EM subsystem development program included thruster optimization, power-supply circuit optimization and flight packaging, subsystem integration, and subsystem acceptance testing including a cyclic test of the total propulsion package.

  2. Electron Transport in Hall Thrusters

    NASA Astrophysics Data System (ADS)

    McDonald, Michael Sean

    Despite high technological maturity and a long flight heritage, computer models of Hall thrusters remain dependent on empirical inputs and a large part of thruster development to date has been heavily experimental in nature. This empirical approach will become increasingly unsustainable as new high-power thrusters tax existing ground test facilities and more exotic thruster designs stretch and strain the boundaries of existing design experience. The fundamental obstacle preventing predictive modeling of Hall thruster plasma properties and channel erosion is the lack of a first-principles description of electron transport across the strong magnetic fields between the cathode and anode. In spite of an abundance of proposed transport mechanisms, accurate assessments of the magnitude of electron current due to any one mechanism are scarce, and comparative studies of their relative influence on a single thruster platform simply do not exist. Lacking a clear idea of what mechanism(s) are primarily responsible for transport, it is understandably difficult for the electric propulsion scientist to focus his or her theoretical and computational tools on the right targets. This work presents a primarily experimental investigation of collisional and turbulent Hall thruster electron transport mechanisms. High-speed imaging of the thruster discharge channel at tens of thousands of frames per second reveals omnipresent rotating regions of elevated light emission, identified with a rotating spoke instability. This turbulent instability has been shown through construction of an azimuthally segmented anode to drive significant cross-field electron current in the discharge channel, and suggestive evidence points to its spatial extent into the thruster near-field plume as well. Electron trajectory simulations in experimentally measured thruster electromagnetic fields indicate that binary collisional transport mechanisms are not significant in the thruster plume, and experiments

  3. The MPD thruster program at JPL

    NASA Technical Reports Server (NTRS)

    Barnett, John; Goodfellow, Keith; Polk, James; Pivirotto, Thomas

    1991-01-01

    The main topics covered include: (1) the Space Exploration Initiative (SEI) context; (2) critical issues of MPD Thruster design; and (3) the Magnetoplasmadynamic (MPD) Thruster Program at JPL. Under the section on the SEI context the nuclear electric propulsion system and some electric thruster options are addressed. The critical issues of MPD Thruster development deal with the requirements, status, and approach taken. The following areas are covered with respect to the MPD Thruster Program at JPL: (1) the radiation-cooled MPD thruster; (2) the High-Current Cathode Test Facility; (3) thruster component thermal modeling; and (4) alkali metal propellant studies.

  4. Laboratory Model 50 kW Hall Thruster

    NASA Technical Reports Server (NTRS)

    Manzella, David; Jankovsky, Robert; Hofer, Richard

    2002-01-01

    A 0.46 meter diameter Hall thruster was fabricated and performance tested at powers up to 72 kilowatts. Thrusts up to 2.9 Newtons were measured. Discharge specific impulses ranged from 1750 to 3250 seconds with discharge efficiencies between 46 and 65 percent. Overall specific impulses ranged from 1550 to 3050 seconds with overall efficiencies between 40 and 57 percent. Performance data indicated significant fraction of multiple-charged ions during operation at elevated power levels. Cathode mass flow rate was shown to be a significant parameter with regard to thruster efficiency.

  5. Magnetic Field Tailored Annular Hall Thruster with Anode Layer

    NASA Astrophysics Data System (ADS)

    Lee, Seunghun; Kim, Holak; Kim, Junbum; Lim, Youbong; Choe, Wonho; Korea Adv Inst of Sci; Tech Team; Korea Institute of Materials Science Collaboration

    2016-09-01

    Plasma propulsion system is one of the key components for advanced missions of satellites as well as deep space exploration. A typical plasma propulsion system is Hall effect thruster that uses crossed electric and magnetic fields to ionize a propellant gas and to accelerate the ionized gas to generate momentum. In Hall thruster plasmas, magnetic field configuration is important due to the fact that electron confinement in the electromagnetic fields affects both plasma and ion beam characteristics as well as thruster performance parameters including thrust, specific impulse, power efficiency, and life time. In this work, development of an anode layer Hall thruster (TAL) with magnetic field tailoring has been attempted. The TAL is possible to keep discharge in 1 to 2 kilovolts of anode voltage, which is useful to obtain high specific impulse. The magnetic field tailoring is used to minimize undesirable heat dissipation and secondary electron emission from the wall surrounding the plasma. We will report 3 W and 200 W thrusters performances measured by a pendulum thrust stand according to the magnetic field configuration. Also, the measured result will be compared with the plasma diagnostics conducted by an angular Faraday probe, a retarding potential analyzer, and a ExB probe.

  6. Performance of a Permanent-Magnet Cylindrical Hall-Effect Thruster

    NASA Technical Reports Server (NTRS)

    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.

  7. Contamination control and plume assessment of low-energy thrusters

    NASA Technical Reports Server (NTRS)

    Scialdone, John J.

    1993-01-01

    Potential contamination of a spacecraft cryogenic surface by a xenon (Xe) ion generator was evaluated. The analysis involves the description of the plume exhausted from the generator with its relative component fluxes on the spacecraft surfaces, and verification of the conditions for condensation, adsorption, and sputtering at those locations. The data describing the plume fluxes and their effects on surfaces were obtained from two sources: the tests carried out with the Xe generator in a small vacuum chamber to indicate deposits and sputter on monitor slides; and the extensive tests with a mercury (Hg) ion thruster in a large vacuum chamber. The Hg thruster tests provided data on the neutrals, on low-energy ion fluxes, on high-energy ion fluxes, and on sputtered materials at several locations within the plume.

  8. Inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1978-01-01

    Inert gas thrusters have continued to be of interest for space propulsion applications. Xenon is of interest in that its physical characteristics are well suited to propulsion. High atomic weight and low tankage fraction were major factors in this choice. If a large amount of propellant was required, so that cryogenic storage was practical, argon is a more economical alternative. Argon was also the preferred propellant for ground applications of thruster technology, such as sputter etching and deposition. Additional magnetic field measurements are reported. These measurements should be of use in magnetic field design. The diffusion of electrons through the magnetic field above multipole anodes was studied in detail. The data were consistent with Bohm diffusion across a magnetic field. The theory based on Bohm diffusion was simple and easily used for diffusion calculations. Limited startup data were obtained for multipole discharge chambers. These data were obtained with refractory cathodes, but should be useful in predicting the upper limits for starting with hollow cathodes.

  9. Large inert-gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Robinson, R. S.

    1981-01-01

    Using present technology as a starting point, performance predictions were made for large thrusters. The optimum beam diameter for maximum thruster efficiency was determined for a range of specific impulse. This optimum beam diameter varied greatly with specific impulse, from about 0.6 m at 3000 seconds (and below) to about 4 m at 10,000 seconds with argon, and from about 0.6 m at 2,000 seconds (and below) to about 12 m at 10,000 seconds with Xe. These beams sizes would require much larger thrusters than those presently available, but would offer substantial complexity and cost reductions for large electric propulsion systems.

  10. Temperature Gradient in Hall Thrusters

    SciTech Connect

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

    2003-11-24

    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.

  11. Azimuthal Spoke Propagation in Hall Effect Thrusters

    NASA Technical Reports Server (NTRS)

    Sekerak, Michael J.; Longmier, Benjamin W.; Gallimore, Alec D.; Brown, Daniel L.; Hofer, Richard R.; Polk, James E.

    2013-01-01

    Spokes are azimuthally propagating perturbations in the plasma discharge of Hall Effect Thrusters (HETs) that travel in the E x B direction and have been observed in many different systems. The propagation of azimuthal spokes are investigated in a 6 kW HET known as the H6 using ultra-fast imaging and azimuthally spaced probes. A spoke surface is a 2-D plot of azimuthal light intensity evolution over time calculated from 87,500 frames/s videos. The spoke velocity has been determined using three methods with similar results: manual fitting of diagonal lines on the spoke surface, linear cross-correlation between azimuthal locations and an approximated dispersion relation. The spoke velocity for three discharge voltages (300, 400 and 450 V) and three anode mass flow rates (14.7, 19.5 and 25.2 mg/s) yielded spoke velocities between 1500 and 2200 m/s across a range of normalized magnetic field settings. The spoke velocity was inversely dependent on magnetic field strength for low B-field settings and asymptoted at B-field higher values. The velocities and frequencies are compared to standard drifts and plasma waves such as E x B drift, electrostatic ion cyclotron, magnetosonic and various drift waves. The empirically approximated dispersion relation yielded a characteristic velocity that matched the ion acoustic speed for 5 eV electrons that exist in the near-anode and near-field plume regions of the discharge channel based on internal measurements. Thruster performance has been linked to operating mode where thrust-to-power is maximized when azimuthal spokes are present so investigating the underlying mechanism of spokes will benefit thruster operation.

  12. Hall thruster with grooved walls

    NASA Astrophysics Data System (ADS)

    Li, Hong; Ning, Zhongxi; Yu, Daren

    2013-02-01

    Axial-oriented and azimuthal-distributed grooves are formed on channel walls of a Hall thruster after the engine undergoes a long-term operation. Existing studies have demonstrated the relation between the grooves and the near-wall physics, such as sheath and electron near-wall transport. The idea to optimize the thruster performance with such grooves was also proposed. Therefore, this paper is devoted to explore the effects of wall grooves on the discharge characteristics of a Hall thruster. With experimental measurements, the variations on electron conductivity, ionization distribution, and integrated performance are obtained. The involved physical mechanisms are then analyzed and discussed. The findings help to not only better understand the working principle of Hall thruster discharge but also establish a physical fundamental for the subsequent optimization with artificial grooves.

  13. Plume characteristics of MPD thrusters: A preliminary examination

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.

    1989-01-01

    A diagnostics facility for MPD thruster plume measurements was built and is currently undergoing testing. The facility includes electrostatic probes for electron temperature and density measurements, Hall probes for magnetic field and current distribution mapping, and an imaging system to establish the global distribution of plasma species. Preliminary results for MPD thrusters operated at power levels between 30 and 60 kW with solenoidal applied magnetic fields show that the electron density decreases exponentially from 1x10(2) to 2x10(18)/cu m over the first 30 cm of the expansion, while the electron temperature distribution is relatively uniform, decreasing from approximately 2.5 eV to 1.5 eV over the same distance. The radiant intensity of the ArII 4879 A line emission also decays exponentially. Current distribution measurements indicate that a significant fraction of the discharge current is blown into the plume region, and that its distribution depends on the magnitudes of both the discharge current and the applied magnetic field.

  14. Performance of a quasi-steady, multi megawatt, coaxial plasma thruster

    NASA Technical Reports Server (NTRS)

    Scheuer, Jay T.; Schoenberg, Kurt F.; Henins, Ivars; Gerwin, Richard A.; Moses, Ronald W., Jr.; Garcia, Jose A.; Gribble, Robert F.; Hoyt, Robert P.; Black, Dorwin C.; Mayo, Robert M.

    1994-01-01

    The Los Alamos National Laboratory Coaxial Thruster Experiment (CTX) has been upgraded to enable the quasisteady operation of magnetoplasmadynamic (MPD) type thrusters at power levels from 2 to 40 MW for 10 ms. Diagnostics include an eight position, three axis magnetic field probe to measure magnetic field fluctuations during the pulse; a triple Langmuir probe to measure ion density, electron temperature, and plasma potential; and a time-of-flight neutral particle spectrometer to measure specific impulse. Here we report on the experimental observations and associated analysis and interpretation of long-pulse, quasisteady, coaxial thruster performance in the CTX device.

  15. Empirical electron cross-field mobility in a Hall effect thruster

    SciTech Connect

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

    2009-10-05

    Electron transport across the magnetic field in Hall effect thrusters is still an open question. Models have so far assumed 1/B{sup 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{sup 2} or 1/B mobility laws for discharge voltages in the 500-700 V range.

  16. Development of a direct evaporation bismuth Hall thruster

    NASA Astrophysics Data System (ADS)

    Massey, Dean Richard

    diffusion bonding process was also developed to join the molybdenum porous disc to the molybdenum anode. Operation of the direct evaporation bismuth Hall thruster revealed interesting phenomenon. By utilizing constant current mode on a discharge power supply, the discharge voltage settles out to a stable operating point which is a function of discharge current, anode face area and average pore size on the vaporizer. Oscillations with a 40 second period were also observed. Preliminary performance data suggests that the direct evaporation bismuth Hall thruster performs similar to xenon and krypton Hall thrusters. Plume interrogation with a Retarding Potential Analyzer confirmed that bismuth ions were being efficiently accelerated while Faraday probe data gave a view of the ion density in the exhausted plume.

  17. Simulation of Electric Propulsion Thrusters

    DTIC Science & Technology

    2011-01-01

    Electric Propulsion Thrusters RTO-EN-AVT-194 17 - 3 [3] is a well-developed, highly successful numerical technique for simulating rarefied gas flows ... Rarefied Flows (Modeles et methodes de calcul des coulements de gaz rarefies ). RTO-EN-AVT-194 14. ABSTRACT Electric propulsion thrusters are replacing...METHODS The focus of this article is on numerical methods used to model the flow of gas and plasma through electric propulsion devices. Discussion

  18. Superconducting electromagnetic thruster

    SciTech Connect

    Meng, J.

    1993-02-11

    An electromagnetic thruster for marine vehicles using a jet of water driven by the interaction of a mutually perpendicular intensified magnetic field and an intensified electric field is disclosed. The intensified magnetic field is produced by superconducting coils cooled by a coolant such as liquid helium. An intensified electric field is produced by passing high amperage current across the seawater jet. These interacting fields produce a Lorentz force perpendicular to mutually perpendicular electric and magnetic field vectors which is used to drive the seawater jet. In some embodiments, the force may also be used to draw water into the jet from the boundary layer flow around the vehicle thereby reducing boundary layer turbulence and associated radiated noise.

  19. Laser-heated thruster

    NASA Technical Reports Server (NTRS)

    Kemp, N. H.; Krech, R. H.

    1980-01-01

    The development of computer codes for the thrust chamber of a rocket of which the propellant gas is heated by a CW laser beam was investigated. The following results are presented: (1) simplified models of laser heated thrusters for approximate parametric studies and performance mapping; (3) computer programs for thrust chamber design; and (3) shock tube experiment to measure absorption coefficients. Two thrust chamber design programs are outlined: (1) for seeded hydrogen, with both low temperature and high temperature seeds, which absorbs the laser radiation continuously, starting at the inlet gas temperature; and (2) for hydrogen seeded with cesium, in which a laser supported combustion wave stands near the gas inlet, and heats the gas up to a temperature at which the gas can absorb the laser energy.

  20. Laser-heated thruster

    NASA Technical Reports Server (NTRS)

    Kemp, N. H.; Lewis, P. F.

    1980-01-01

    The development of a computer program for the design of the thrust chamber for a CW laser heated thruster was examined. Hydrodgen was employed as the propellant gas and high temperature absorber. The laser absorption coefficient of the mixture/laser radiation combination is given in temperature and species densities. Radiative and absorptive properties are given to determine radiation from such gas mixtures. A computer code for calculating the axisymmetric channel flow of a gas mixture in chemical equilibrium, and laser energy absorption and convective and radiative heating is described. It is concluded that: (1) small amounts of cesium seed substantially increase the absorption coefficient of hydrogen; (2) cesium is a strong radiator and contributes greatly to radiation of cesium seeded hydrogen; (3) water vapor is a poor absorber; and (4) for 5.3mcm radiation, both H2O/CO and NO/CO seeded hydrogen mixtures are good absorbers.

  1. Development And Testing Of The Inertial Electrostatic Confinement Diffusion Thruster

    NASA Technical Reports Server (NTRS)

    Becnel, Mark D.; Polzin, Kurt A.

    2013-01-01

    The Inertial Electrostatic Confinement (IEC) diffusion thruster is an experiment in active development that takes advantage of physical phenomenon that occurs during operation of an IEC device. The IEC device has been proposed as a fusion reactor design that relies on traditional electrostatic ion acceleration and is typically arranged in a spherical geometry. The design incorporates two radially-symmetric spherical electrodes. Often the inner electrode utilizes a grid of wire shaped in a sphere with a radius 15 to 50 percent of the radius of the outer electrode. The inner electrode traditionally has 90 percent or more transparency to allow particles (ions) to pass to the center of the spheres and collide/recombine in the dense plasma core at r=0. When operating the IEC, an unsteady plasma leak is typically observed passing out one of the gaps in the lattice grid of the inner electrode. The IED diffusion thruster is based upon the idea that this plasma leak can be used for propulsive purposes. The IEC diffusion thruster utilizes the radial symmetry found in the IEC device. A cylindrical configuration is employed here as it will produce a dense core of plasma the length of the cylindrical grid while promoting the plasma leak to exhaust through an electromagnetic nozzle at one end of the apparatus. A proof-of-concept IEC diffusion thruster is operational and under testing using argon as propellant (Figure 1).

  2. Deep Space Mission Applications for NEXT: NASA's Evolutionary Xenon Thruster

    NASA Technical Reports Server (NTRS)

    Oh, David; Benson, Scott; Witzberger, Kevin; Cupples, Michael

    2004-01-01

    NASA's Evolutionary Xenon Thruster (NEXT) is designed to address a need for advanced ion propulsion systems on certain future NASA deep space missions. This paper surveys seven potential missions that have been identified as being able to take advantage of the unique capabilities of NEXT. Two conceptual missions to Titan and Neptune are analyzed, and it is shown that ion thrusters could decrease launch mass and shorten trip time, to Titan compared to chemical propulsion. A potential Mars Sample return mission is described, and compassion made between a chemical mission and a NEXT based mission. Four possible near term applications to New Frontiers and Discovery class missions are described, and comparisons are made to chemical systems or existing NSTAR ion propulsion system performance. The results show that NEXT has potential performance and cost benefits for missions in the Discovery, New Frontiers, and larger mission classes.

  3. Wall current closure effects on plasma and sheath fluctuations in Hall thrusters

    SciTech Connect

    Frias, Winston Smolyakov, Andrei I.; Kaganovich, Igor D.; Raitses, Yevgeny

    2014-06-15

    The excitation of negative energy, ion sound type modes driven by the E × B drift and the reactive/dissipative response of the wall sheath interface is analyzed for conditions typical in a Hall thruster. Such sheath impedance modes are sensitive to the dielectric properties of the thruster wall material, which therefore may have direct influence (other than via the secondary electron emission) on fluctuations and transport. Our results predict mode frequencies consistent with the frequencies of fluctuations observed experimentally.

  4. Erosion Measurements in a Diverging Cusped-Field Thruster (Pre Print)

    DTIC Science & Technology

    2012-02-01

    Ion Velocity Measurements Within the Acceleration Channel of a Low-Power Hall Thruster,” Plasma Science, IEEE Transactions on, Vol. 36, No. 5...Measurements Within the Acceleration Channel of a Low-Power Hall Thruster,” Plasma Science, IEEE Transactions on, Vol. 38, No. 4, April 2010, pp. 1052...etry,” Physics of Plasmas , Vol. 18, 2011, pp. 033509. 62 Dunaevsky, A., Raitses, Y., and Fisch, N. J., “Secondary Electron Emission from Dielectric

  5. Challenges to computing plasma thruster dynamics

    SciTech Connect

    Smith, G.A. )

    1992-01-01

    This paper describes computational challenges in describing high thrust and I[sub sp] expected from the proposed ion-compressed antimatter nuclear (ICAN) propulsion system. This concept uses antiprotons to induce fission reactions that jump start a microfission/fusion process in a target compressed by low-energy ion beams. The ICAN system could readily provide the high energy density required for interplanetary space missions of short duration. In conventional rocket design, thrust is obtained by expelling a propellant under high pressure through a nozzle. A larger I[sub sp] can be achieved by operating the system at a higher temperature. Full ionization of propellant at high temperature introduces new and challenging questions in the design of plasma thrusters.

  6. High-Power Hall Thruster Technology Evaluated for Primary Propulsion Applications

    NASA Technical Reports Server (NTRS)

    Manzella, David H.; Jankovsky, Robert S.; Hofer, Richard R.

    2003-01-01

    High-power electric propulsion systems have been shown to be enabling for a number of NASA concepts, including piloted missions to Mars and Earth-orbiting solar electric power generation for terrestrial use (refs. 1 and 2). These types of missions require moderate transfer times and sizable thrust levels, resulting in an optimized propulsion system with greater specific impulse than conventional chemical systems and greater thrust than ion thruster systems. Hall thruster technology will offer a favorable combination of performance, reliability, and lifetime for such applications if input power can be scaled by more than an order of magnitude from the kilowatt level of the current state-of-the-art systems. As a result, the NASA Glenn Research Center conducted strategic technology research and development into high-power Hall thruster technology. During program year 2002, an in-house fabricated thruster, designated the NASA-457M, was experimentally evaluated at input powers up to 72 kW. These tests demonstrated the efficacy of scaling Hall thrusters to high power suitable for a range of future missions. Thrust up to nearly 3 N was measured. Discharge specific impulses ranged from 1750 to 3250 sec, with discharge efficiencies between 46 and 65 percent. This thruster is the highest power, highest thrust Hall thruster ever tested.

  7. Study of Breathing Oscillations in a Hall Thruster

    NASA Astrophysics Data System (ADS)

    Keller, Scott; Raitses, Yevgeny; Diallo, Ahmed

    2014-10-01

    Breathing oscillations are the most powerful low frequency (10-30 kHz) oscillations that are typically observed in different types of Hall thrusters. We report on investigations of the effects of both natural and artificially driven breathing oscillations on the discharge and plasma properties of a cylindrical Hall thruster. In order to produce artificially coherent oscillations, a sinusoidal modulation up to 30 VAC of the anode potential in the range of 5-30 kHz is applied to the thruster. These driven modes are studied in operating regimes with and without naturally occurring oscillations. The imposed periodicity allows for measurement of the time-dependent ion velocity distribution through a novel heterodyne approach to laser-induced fluorescence (LIF) using phase-sensitive detection. Further comparison between natural and driven modes is performed through the analysis of the discharge and ion currents, as well as high-speed imaging data. Results serve both to validate the LIF technique and to improve understanding of breathing oscillations. In particular, we show oscillations of the ion velocity distribution function due to breathing oscillations and explain their correlation with oscillations of the discharge and ion currents. This work was supported by DOE Contract DE-AC02-09CH11466.

  8. Direct Drive Hall Thruster System Development

    NASA Technical Reports Server (NTRS)

    Hoskins, W. Andrew; Homiak, Daniel; Cassady, R. Joseph; Kerslake, Tom; Peterson, Todd; Ferguson, Dale; Snyder, Dave; Mikellides, Ioannis; Jongeward, Gary; Schneider, Todd

    2003-01-01

    The sta:us of development of a Direct Drive Ha!! Thruster System is presented. 13 the first part. a s:udy of the impacts to spacecraft systems and mass benefits of a direct-drive architecture is reviewed. The study initially examines four cases of SPT-100 and BPT-4000 Hall thrusters used for north-south station keeping on an EXPRESS-like geosynchronous spacecraft and for primary propulsion for a Deep Space- 1 based science spacecraft. The study is also extended the impact of direct drive on orbit raising for higher power geosynchronous spacecraft and on other deep space missions as a function of power and delta velocity. The major system considerations for accommodating a direct drive Hall thruster are discussed, including array regulation, system grounding, distribution of power to the spacecraft bus, and interactions between current-voltage characteristics for the arrays and thrusters. The mass benefit analysis shows that, for the initial cases, up to 42 kg of dry mass savings is attributable directly to changes in the propulsion hardware. When projected mass impacts of operating the arrays and the electric power system at 300V are included, up to 63 kg is saved for the four initial cases. Adoption of high voltage lithium ion battery technology is projected to further improve these savings. Orbit raising of higher powered geosynchronous spacecraft, is the mission for which direct drive provides the most benefit, allowing higher efficiency electric orbit raising to be accomplished in a limited period of time, as well as nearly eliminating significant power processing heat rejection mass. The total increase in useful payload to orbit ranges up to 278 kg for a 25 kW spacecraft, launched from an Atlas IIA. For deep space missions, direct drive is found to be most applicable to higher power missions with delta velocities up to several km/s , typical of several Discovery-class missions. In the second part, the status of development of direct drive propulsion power

  9. Inert gas thrusters

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.

    1977-01-01

    Inert gases, particularly argon and xenon, are of interest as possible alternatives to the usual electric thruster propellants of mercury and cesium. Hollow cathode data were obtained for a wide range of operating conditions. Some test conditions gave plasma coupling voltages at or below the sputtering threshold, hence should permit long operating lifetimes. All observations of hollow cathode operation were consistent with a single theory of operation, in which a significant amount of the total electron emission is from localized areas within the orifice. This mode of emission is also supported by scanning electron microscope photographs that indicate local temperatures at or near the melting temperature of the tungsten tip. Experimental hollow cathode performance was correlated for two orifice diameters, three inert gas propellants, and a range of flow rates for each propellant. The basic theory for the production of doubly ionized argon and xenon was completed. Experimental measurements of the doubly ionized fraction agree with theory within about plus or minus 20 percent. High voltage isolators were studied for the propellant feed line. The breakdown voltage per segment ranged from 300 to over 500 V with argon.

  10. Magnetic shielding of a laboratory Hall thruster. II. Experiments

    SciTech Connect

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

    2014-01-28

    The physics of magnetic shielding in Hall thrusters were validated through laboratory experiments demonstrating essentially erosionless, high-performance operation. The magnetic field near the walls of a laboratory Hall thruster was modified to effectively eliminate wall erosion while maintaining the magnetic field topology away from the walls necessary to retain efficient operation. Plasma measurements at the walls validate our understanding of magnetic shielding as derived from the theory. The plasma potential was maintained very near the anode potential, the electron temperature was reduced by a factor of two to three, and the ion current density was reduced by at least a factor of two. Measurements of the carbon backsputter rate, wall geometry, and direct measurement of plasma properties at the wall indicate that the wall erosion rate was reduced by a factor of 1000 relative to the unshielded thruster. These changes effectively eliminate wall erosion as a life limitation in Hall thrusters, enabling a new class of deep-space missions that could not previously be attempted.

  11. Helicon plasma thruster discharge model

    SciTech Connect

    Lafleur, T.

    2014-04-15

    By considering particle, momentum, and energy balance equations, we develop a semi-empirical quasi one-dimensional analytical discharge model of radio-frequency and helicon plasma thrusters. The model, which includes both the upstream plasma source region as well as the downstream diverging magnetic nozzle region, is compared with experimental measurements and confirms current performance levels. Analysis of the discharge model identifies plasma power losses on the radial and back wall of the thruster as the major performance reduction factors. These losses serve as sinks for the input power which do not contribute to the thrust, and which reduce the maximum plasma density and hence propellant utilization. With significant radial plasma losses eliminated, the discharge model (with argon) predicts specific impulses in excess of 3000 s, propellant utilizations above 90%, and thruster efficiencies of about 30%.

  12. Effect of azimuthal diversion rail on an ATON-type Hall thruster

    NASA Astrophysics Data System (ADS)

    Xu, Zhang; Liqiu, Wei; Liang, Han; Yongjie, Ding; Daren, Yu

    2017-03-01

    A newly designed azimuthal diversion rail (ADR) is studied and used to enhance the ionization process in an ATON-type Hall thruster. The diversion rail efficiently reduces the neutral flow axial velocity, and hence, increases the resistance time of atoms in the discharge channel of the Hall thruster. Thrust performances, in terms of thrust, anode efficiency and ion beam divergence, are found to be improved because of the application of the diversion rail, especially at low mass flow rate conditions. Experiment results reveal that the ADR increases the mass utilization under insufficient mass flow rate operating conditions. The design of the ADR broadens the efficient operating range of Hall thrusters and has significant contribution to multi-mode Hall thruster development.

  13. Hall Thruster Plume Measurements On-Board the Russian Express Satellites

    NASA Technical Reports Server (NTRS)

    Manzella, David; Jankovsky, Robert; Elliott, Frederick; Mikellides, Ioannis; Jongeward, Gary; Allen, Doug

    2001-01-01

    The operation of North-South and East-West station-keeping Hall thruster propulsion systems on-board two Russian Express-A geosynchronous communication satellites were investigated through a collaborative effort with the manufacturer of the spacecraft. Over 435 firings of 16 different thrusters with a cumulative run time of over 550 hr were reported with no thruster failures. Momentum transfer due to plume impingement was evaluated based on reductions in the effective thrust of the SPT-100 thrusters and induced disturbance torques determined based on attitude control system data and range data. Hall thruster plasma plume effects on the transmission of C-band and Ku-band communication signals were shown to be negligible. On-orbit ion current density measurements were made and subsequently compared to predictions and ground test data. Ion energy, total pressure, and electric field strength measurements were also measured on-orbit. The effect of Hall thruster operation on solar array performance over several months was investigated. A subset of these data is presented.

  14. NASA GRC High Power Electromagnetic Thruster Program

    NASA Astrophysics Data System (ADS)

    Lapointe, Michael R.; Pencil, Eric J.

    2004-02-01

    Interest in high power electromagnetic propulsion has been revived to support a variety of future space missions, such as platform maneuvering in low earth orbit, cost-effective cargo transport to lunar and Mars bases, asteroid and outer planet sample return, deep space robotic exploration, and piloted missions to Mars and the outer planets. Magnetoplasmadynamic (MPD) thrusters have demonstrated, at the laboratory level, the capacity to process megawatts of electrical power while providing higher thrust densities than current electric propulsion systems. The ability to generate higher thrust densities permits a reduction in the number of thrusters required to perform a given mission and alleviates the system complexity associated with multiple thruster arrays. The specific impulse of an MPD thruster can be optimized to meet given mission requirements, from a few thousand seconds with heavier gas propellants up to 10,000 seconds with hydrogen propellant. In support of NASA space science and human exploration strategic initiatives, Glenn Research Center is developing and testing pulsed, MW-class MPD thrusters as a prelude to long-duration high power thruster tests. The research effort includes numerical modeling of self-field and applied-field MPD thrusters and experimental testing of quasi-steady MW-class MPD thrusters in a high power pulsed thruster facility. This paper provides an overview of the GRC high power electromagnetic thruster program and the pulsed thruster test facility.

  15. Scale Model Thruster Acoustic Measurement Results

    NASA Technical Reports Server (NTRS)

    Kenny, R. Jeremy; Vargas, Magda B.

    2013-01-01

    Subscale rocket acoustic data is used to predict acoustic environments for full scale rockets. Over the last several years acoustic data has been collected during horizontal tests of solid rocket motors. Space Launch System (SLS) Scale Model Acoustic Test (SMAT) was designed to evaluate the acoustics of the SLS vehicle including the liquid engines and solid rocket boosters. SMAT is comprised of liquid thrusters scalable to the Space Shuttle Main engines (SSME) and Rocket Assisted Take Off (RATO) motors scalable to the 5-segment Reusable Solid Rocket Motor (RSTMV). Horizontal testing of the liquid thrusters provided an opportunity to collect acoustic data from liquid thrusters to characterize the acoustic environments. Acoustic data was collected during the horizontal firings of a single thruster and a 4-thruster (Quad) configuration. Presentation scope. Discuss the results of the single and 4-thruster acoustic measurements. Compare the measured acoustic levels of the liquid thrusters to the Solid Rocket Test Motor V - Nozzle 2 (SRTMV-N2).

  16. A Small Modular Laboratory Hall Effect Thruster

    NASA Astrophysics Data System (ADS)

    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.

  17. Performance characteristics according to the channel length and magnetic fields of cylindrical Hall thrusters

    NASA Astrophysics Data System (ADS)

    Lee, Jongsub; Seo, Mihui; Seon, Jongho; June Lee, Hae; Choe, Wonho

    2011-09-01

    Performance characteristics of low power cylindrical Hall thrusters are investigated in terms of the length of the discharge channel. Thrust, efficiency, discharge current, and propellant utilization are evaluated for different channel lengths of 19, 22, and 25 mm. It is found that the propellant utilization and ion energy distribution function are strongly associated with the channel length. Increase of thrust and efficiency are also found with increasing channel lengths. These characteristics of the thruster are interpreted with possible generation of multi-charged ions due to increased residing time within the extended space inside the channel.

  18. Design of an ICRF plasma thruster antenna by TOPICA

    NASA Astrophysics Data System (ADS)

    Vecchi, Giuseppe; Lancellotti, Vito; Maggiora, Riccardo

    2006-10-01

    A typical RF plasma thruster is comprised of an RF plasma source, an open-ended magnetic confinement device, an RF acceleration unit and a magnetic nozzle. The usual choice for the acceleration is to employ the Ion-Cyclotron resonance frequency (ICRF), a well established technology in fusion experiments for transferring large RF powers to magnetized plasmas. To help design RF thruster ICRF antennas, TOPICA (Torino Polytechnic Ion Cyclotron Antenna) code [1] has been recently extended to handle cylindrically symmetric plasmas. The latter entailed developing a wholly new module of TOPICA charged with the task of solving Maxwell's equations in cylindrical magnetized warm plasmas and yielding the Green's functionY (m,kz), i.e. the relationship at the air-plasma interface between the transverse magnetic and electric fields in the spectral (wavenumber) domain. The approach to the problem of determining the antenna input impedance relies on an integral-equation formulation for the self-consistent evaluation of the current distribution on the conductors. This work reports on TOPICA evolution and presents the design of an RF thruster ICRF antenna. *V. Lancellotti et al., Nucl. Fusion, 46 (2006) S476-S499

  19. Krypton charge exchange cross sections for Hall effect thruster models

    SciTech Connect

    Hause, Michael L.; Prince, Benjamin D.; Bemish, Raymond J.

    2013-04-28

    Following discharge from a Hall effect thruster, charge exchange occurs between ions and un-ionized propellant atoms. The low-energy cations produced can disturb operation of onboard instrumentation or the thruster itself. Charge-exchange cross sections for both singly and doubly charged propellant atoms are required to model these interactions. While xenon is the most common propellant currently used in Hall effect thrusters, other propellants are being considered, in particular, krypton. We present here guided-ion beam measurements and comparisons to semiclassical calculations for Kr{sup +} + Kr and Kr{sup 2+} + Kr cross sections. The measurements of symmetric Kr{sup +} + Kr charge exchange are in good agreement with both the calculations including spin-orbit effects and previous measurements. For the symmetric Kr{sup 2+} + Kr reaction, we present cross section measurements for center-of-mass energies between 1 eV and 300 eV, which spans energies not previously examined experimentally. These cross section measurements compare well with a simple one-electron transfer model. Finally, cross sections for the asymmetric Kr{sup 2+} + Kr {yields} Kr{sup +} + Kr{sup +} reaction show an onset near 12 eV, reaching cross sections near constant value of 1.6 A{sup 2} with an exception near 70-80 eV.

  20. An Overview of the VHITAL Program: A Two-Stage Bismuth Fed Very High Specific Impulse Thruster with Anode Layer

    NASA Technical Reports Server (NTRS)

    Sengupta, Anita; Marrese-Reading, Colleen; Capelli, Mark; Scharfe, David; Tverdokhlebov, Sergey; Semenkin, Sasha; Tverdokhlebov, Oleg; Boyd, Ian; Keidar, Michael; Yalin, Azer; Markusic, Tom; Polzin, Kurt

    2005-01-01

    The Very High Isp Thruster with Anode Layer (VHITAL) is a two stage Hall thruster program that is a part of NASA's Prometheus Program in NASA's New Exploration Systems Mission Directorate (ESMD). It is a potentially viable low-cost alternative to ion engines for near-term NEP applications with the growth potential to support mid-term and far-term NEP missions... This paper will present an overview of the thruster fabrication, pre-existing TAL 160 demonstration, feed system development, lifetime assessment, contamination assessment, and mission study activities performed to date.