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Sample records for acceleration space thruster

  1. The FAST (FRC Acceleration Space Thruster) Experiment

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  2. The FRC Acceleration Space Thruster (FAST) Experiment

    NASA Technical Reports Server (NTRS)

    Martin, Adam; Eskridge, Richard; Houts, Mike; Slough, John; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    The objective of the FRC (Field Reversed Configuration) Acceleration Space Thruster (FAST) Experiment is to investigate the use of a repetitive FRC source as a thruster, specifically for an NEP (nuclear electric propulsion) system. The Field Reversed Configuration is a plasmoid with a closed poloidal field line structure, and has been extensively studied as a fusion reactor core. An FRC thruster works by repetitively producing FRCs and accelerating them to high velocity. An FRC thruster should be capable of I(sub sp)'s in the range of 5,000 - 25,000 seconds and efficiencies in the range of 60 - 80 %. In addition, they can have thrust densities as high as 10(exp 6) N/m2, and as they are inductively formed, they do not suffer from electrode erosion. The jet-power should be scalable from the low to the high power regime. The FAST experiment consists of a theta-pinch formation chamber, followed by an acceleration stage. Initially, we will produce and accelerate single FRCs. The initial focus of the experiment will be on the ionization, formation and acceleration of a single plasmoid, so as to determine the likely efficiency and I(sub sp). Subsequently, we will modify the device for repetitive burst-mode operation (5-10 shots). A variety of diagnostics are or will be available for this work, including a HeNe interferometer, high-speed cameras, and a Thomson-scattering system. The status of the experiment will be described.

  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. Miniature Free-Space Electrostatic Ion Thrusters

    NASA Technical Reports Server (NTRS)

    Hartley, Frank T.; Stephens, James B.

    2006-01-01

    A miniature electrostatic ion thruster is proposed for maneuvering small spacecraft. In a thruster based on this concept, one or more propellant gases would be introduced into an ionizer based on the same principles as those of the device described in an earlier article, "Miniature Bipolar Electrostatic Ion Thruster". On the front side, positive ions leaving an ionizer element would be accelerated to high momentum by an electric field between the ionizer and an accelerator grid around the periphery of the concave laminate structure. On the front side, electrons leaving an ionizer element 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 a thruster design consisting of multiple membrane ionizers in a thin laminate structure with a peripheral accelerator grid, the direction of thrust could then be controlled (without need for moving parts in the thruster) by regulating the supply of gas to specific ionizer.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. 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.

  10. NASA - 77M prototype hall thruster built under the High Voltage Hall accelerator development project

    NASA Technical Reports Server (NTRS)

    2005-01-01

    NASA - 77M prototype hall thruster built under the High Voltage Hall accelerator development project funded by the Science Mission Directorate ; potential use is propulsion for deep space science missions

  11. 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.

  12. Plasma acceleration processes in an ablative pulsed plasma thruster

    SciTech Connect

    Koizumi, Hiroyuki; Noji, Ryosuke; Komurasaki, Kimiya; Arakawa, Yoshihiro

    2007-03-15

    Plasma acceleration processes in an ablative pulsed plasma thruster (APPT) were investigated. APPTs are space propulsion options suitable for microspacecraft, and have recently attracted much attention because of their low electric power requirements and simple, compact propellant system. The plasma acceleration mechanism, however, has not been well understood. In the present work, emission spectroscopy, high speed photography, and magnetic field measurements are conducted inside the electrode channel of an APPT with rectangular geometry. The successive images of neutral particles and ions give us a comprehensive understanding of their behavior under electromagnetic acceleration. The magnetic field profile clarifies the location where the electromagnetic force takes effect. As a result, it is shown that high density, ablated neutral gas stays near the propellant surface, and only a fraction of the neutrals is converted into plasma and electromagnetically accelerated, leaving the residual neutrals behind.

  13. Acceleration Mechanism Of Pulsed Laser-Electromagnetic Hybrid Thruster

    NASA Astrophysics Data System (ADS)

    Horisawa, Hideyuki; Mashima, Yuki; Yamada, Osamu

    2011-11-01

    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.

  14. A Plasmoid Thruster for Space Propulsion

    NASA Technical Reports Server (NTRS)

    Koelfgen, Syri J.; Hawk, Clark W.; Eskridge, Richard; Smith, James W.; Martin, Adam K.

    2003-01-01

    There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are classified according to the relative strength of the poloidal and toroidal magnetic field (B(sub p), and B(sub t), respectively). An object with B(sub p), / B(sub t) much greater than 1 is classified as a Field Reversed Configuration (FRC); if B(sub p) approximately equal to B(sub t), it is called a Spheromak. The plasmoid thruster operates by producing FRC-like plasmoids and subsequently ejecting them from the device at a high velocity. The plasmoid is formed inside of a single-turn conical theta-pinch coil. As this process is inductive, there are no electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s, and calculations indicate that velocities in excess of 100 km/s should be possible. This concept should be capable of producing Isp's in the range of 5,000 - 15,000 s with thrust densities on the order of 10(exp 5) N per square meters. The current experiment is designed to produce jet powers in the range of 5 - 10 kW, although the concept should be scalable to several MW's. The plasmoid mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras and a laser interferometer. Also of key importance will be measurements of the efficiency and mass utilization. Simulations of the plasmoid thruster using MOQUI, a time-dependent MHD code, will be carried out concurrently with experimental testing.

  15. A Plasmoid Thruster for Space Propulsion

    NASA Technical Reports Server (NTRS)

    Koelfgen, Syri J.; Hawk, Clark W.; Eskridge, Richard; Smith, James W.; Martin, Adam K.

    2003-01-01

    There are a number of possible advantages to using accelerated plasmoids for in-space propulsion. A plasmoid is a compact plasma structure with an integral magnetic field. They have been studied extensively in controlled fusion research and are classified according to the relative strength of the poloidal and toroidal magnetic field (BP and Bt, respectively). An Object with B P t >> 1 is classified as a Field Reverse Configuration (FRC); if B, = Bt, it is called a Spheromak. The plasmoid thruster operates by producing FRC-like plasmoids, and subsequently ejecting them from the device at high velocity. The plasmoid is formed inside of a single turn conical theta-pinch coil. As this process is inductive, there are no electrodes. Similar experiments have yielded plasmoid velocities of at least 50 km/s (l), and calculations indicate that velocities in excess of 100 km/s should be possible. This concept should be capable of producing Isp s in the range of 5,000 - 10,000 s with thrust densities of order 10(exp 5) N/sq m. The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable to several MW s. The plasmoids mass and velocity will be measured with a variety of diagnostics, including internal and external B-dot probes, flux loops, Langmuir probes, high-speed cameras, and a laser interferometer. Also of key importance will be measurements of the efficiency and mass utilization. Simulations of the plasmoid thruster using MOQUI, a time dependent MHD code, will be carried out concurrently with experimental testing.

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

    SciTech Connect

    Takahashi, H.; Yu, A.

    1997-04-01

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

  17. 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.

  18. IEC Thrusters for Space Probe Applications and Propulsion

    SciTech Connect

    Miley, George H.; Momota, Hiromu; Wu Linchun; Reilly, Michael P.; Teofilo, Vince L.; Burton, Rodney; Dell, Richard; Dell, Dick; Hargus, William A.

    2009-03-16

    Earlier conceptual design studies (Bussard, 1990; Miley et al., 1998; Burton et al., 2003) have described Inertial Electrostatic Confinement (IEC) fusion propulsion to provide a high-power density fusion propulsion system capable of aggressive deep space missions. However, this requires large multi-GW thrusters and a long term development program. As a first step towards this goal, a progression of near-term IEC thrusters, stating with a 1-10 kWe electrically-driven IEC jet thruster for satellites are considered here. The initial electrically-powered unit uses a novel multi-jet plasma thruster based on spherical IEC technology with electrical input power from a solar panel. In this spherical configuration, Xe ions are generated and accelerated towards the center of double concentric spherical grids. An electrostatic potential well structure is created in the central region, providing ion trapping. Several enlarged grid opening extract intense quasi-neutral plasma jets. A variable specific impulse in the range of 1000-4000 seconds is achieved by adjusting the grid potential. This design provides high maneuverability for satellite and small space probe operations. The multiple jets, combined with gimbaled auxiliary equipment, provide precision changes in thrust direction. The IEC electrical efficiency can match or exceed efficiencies of conventional Hall Current Thrusters (HCTs) while offering advantages such as reduced grid erosion (long life time), reduced propellant leakage losses (reduced fuel storage), and a very high power-to-weight ratio. The unit is ideally suited for probing missions. The primary propulsive jet enables delicate maneuvering close to an object. Then simply opening a second jet offset 180 degrees from the propulsion one provides a 'plasma analytic probe' for interrogation of the object.

  19. IEC Thrusters for Space Probe Applications and Propulsion

    NASA Astrophysics Data System (ADS)

    Miley, George H.; Momota, Hiromu; Wu, Linchun; Reilly, Michael P.; Teofilo, Vince L.; Burton, Rodney; Dell, Richard; Dell, Dick; Hargus, William A.

    2009-03-01

    Earlier conceptual design studies (Bussard, 1990; Miley et al., 1998; Burton et al., 2003) have described Inertial Electrostatic Confinement (IEC) fusion propulsion to provide a high-power density fusion propulsion system capable of aggressive deep space missions. However, this requires large multi-GW thrusters and a long term development program. As a first step towards this goal, a progression of near-term IEC thrusters, stating with a 1-10 kWe electrically-driven IEC jet thruster for satellites are considered here. The initial electrically-powered unit uses a novel multi-jet plasma thruster based on spherical IEC technology with electrical input power from a solar panel. In this spherical configuration, Xe ions are generated and accelerated towards the center of double concentric spherical grids. An electrostatic potential well structure is created in the central region, providing ion trapping. Several enlarged grid opening extract intense quasi-neutral plasma jets. A variable specific impulse in the range of 1000-4000 seconds is achieved by adjusting the grid potential. This design provides high maneuverability for satellite and small space probe operations. The multiple jets, combined with gimbaled auxiliary equipment, provide precision changes in thrust direction. The IEC electrical efficiency can match or exceed efficiencies of conventional Hall Current Thrusters (HCTs) while offering advantages such as reduced grid erosion (long life time), reduced propellant leakage losses (reduced fuel storage), and a very high power-to-weight ratio. The unit is ideally suited for probing missions. The primary propulsive jet enables delicate maneuvering close to an object. Then simply opening a second jet offset 180 degrees from the propulsion one provides a "plasma analytic probe" for interrogation of the object.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  1. Liquid-Metal-Fed Pulsed Electromagnetic Thrusters For In-Space Propulsion

    NASA Technical Reports Server (NTRS)

    Markusic, T. E.

    2004-01-01

    We describe three pulsed electromagnetic thruster concepts, which span four orders of magnitude in power processing capability (100 W to >100 kW), for in-space propulsion applications. The primary motivation for using a pulsed system is to is to enable high (instantaneous) power operation, which provides high acceleration efficiency, while using considerably less (continuous) power from the spacecraft power system. Unfortunately, conventional pulsed thrusters require failure-prone electrical switches and gas-puff valves. The series of thrusters described here directly address this problem, through the use of liquid metal propellant, by either eliminating both components or providing less taxing operational requirements, thus yielding a path toward both efficient and reliable pulsed electromagnetic thrusters. The emphasis of this paper is to conceptually describe each of the thruster concepts; however, initial test results with gallium propellant in one thruster geometry are presented. These tests reveal that a greater understanding of gallium material compatibility, contamination, and wetting behavior will be necessary before a completely functional thruster can be developed. Initial experimental results aimed at providing insight into these issues are presented.

  2. A novel laser ablation plasma thruster with electromagnetic acceleration

    NASA Astrophysics Data System (ADS)

    Zhang, Yu; Zhang, Daixian; Wu, Jianjun; He, Zhen; Zhang, Hua

    2016-10-01

    A novel laser ablation plasma thruster accelerated by electromagnetic means was proposed and investigated. The discharge characteristics and thrust performance were tested with different charged energy, structural parameters and propellants. The thrust performance was proven to be improved by electromagnetic acceleration. In contrast with the pure laser propulsion mode, the thrust performance in electromagnetic acceleration modes was much better. The effects of electrodes distance and the off-axis distance between ceramic tube and cathode were tested, and it's found that there were optimal structural parameters for achieving optimal thrust performance. It's indicated that the impulse bit and specific impulse increased with increasing charged energy. In our experiments, the thrust performance of the thruster was optimal in large charged energy modes. With the charged energy 25 J and the use of metal aluminum, a maximal impulse bit of 600 μNs, a specific impulse of approximate 8000 s and thrust efficiency of about 90% were obtained. For the PTFE propellant, a maximal impulse bit of about 350 μNs, a specific impulse of about 2400 s, and thrust efficiency of about 16% were obtained. Besides, the metal aluminum was proven to be the better propellant than PTFE for the thruster.

  3. Ion Emission Characteristics of a Forward Laser Accelerated Plasma Thruster

    SciTech Connect

    Oyaizu, Keishi; Izumi, Masaya; Horisawa, Hideyuki; Kimura, Itsuro

    2005-04-27

    A fundamental study of a forward laser accelerated plasma thruster was conducted. In order to evaluate thrust performances of the thruster, a time-of-flight measurement was conducted for an Al-foil target irradiated with an Nd:YAG laser of 1J/pulse with pulse-width of 10nsec. From the measurement, the average plasma speed was about 53 km/sec. Time-gated imaging of the plasma with an ICCD camera was also conducted. From the observation, rapid plasmas were observed on both sides of the target. Each image from the ICCD camera was processed by an image processing software into an emission intensity distribution of the plasma at every 10nsec. Axial velocity of the plasma was estimated from the temporal evolution of the plasma edge. The average and maximum plasma expansion velocities in a forward direction were estimated about 40 km/s and 160 km/sec, respectively.

  4. 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.

  5. 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.

  6. 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.

  7. 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 tested on five different 30-cm diameter bombardment thrustors to evaluate the effects of grid geometry variations on thrustor 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. Also investigated were the effects on discharge chamber performance of main magnetic field changes, magnetic baffle current cathode pole piece length and cathode position.

  8. Development of MPD thruster EM for a space test. [Engineering model

    SciTech Connect

    Shiina, K.; Suzuki, H.; Uematsu, K.; Ohtsuka, T.; Toki, K. Institute of Space and Astronautical Science, Kanagawa )

    1990-01-01

    An engineering model (EM) of MPD thruster has been developed for a space test on board the first Space Flyer Unit (SFU-1). A thermal vacuum test was conducted, and the following results were obtained: (1) a thermal mathematical model of MPD thruster EM was established, (2) sizing data of thruster heaters were obtained, and (3) thermal characteristics of the MPD thruster EM were confirmed to meet the requirement. The data are going to be reflected in designing a protoflight model of MPD thruster. 8 refs.

  9. Characterization of ion accelerating systems on NASA LeRC's ion thrusters

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.

    1992-01-01

    An investigation is conducted regarding ion-accelerating systems for two NASA thrusters to study the limits of ion-extraction capability or perveance. A total of nine two-grid ion-accelerating systems are tested with the 30- and 50-cm-diam ring-cusp inert-gas ion thrusters emphasizing the extension of ion-extraction. The vacuum-tank testing is described using xenon, krypton, and argon propellants, and thruster performance is computed with attention given to theoretical design considerations. Reductions in perveance are noted with decreasing accelerator-hole-to-screen-hole diameter ratios. Perveance values vary indirectly with the ratio of discharge voltage to total accelerating voltage, and screen/accelerator electrode hole-pair alignment is also found to contribute to perveance values.

  10. Space simulation experiments on reaction control system thruster plumes

    NASA Technical Reports Server (NTRS)

    Cassidy, J. F.

    1972-01-01

    A space simulation procedure was developed for studying rocket plume contamination effects using a 5-pound bipropellant reaction control system thruster. Vacuum chamber pressures of 3 x 10 to the minus 5 torr (70 miles altitude) were achieved with the thruster firing in pulse trains consisting of eight pulses (50 msec on, 100 msec off, and seven minutes between pulse trains). The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. An effort was made to simulate propellant system flow dynamics corresponding to actual spacecraft mission use. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.

  11. 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.

  12. 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.

  13. 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.

  14. Electrostatic thrusters.

    NASA Technical Reports Server (NTRS)

    Kaufman, H. R.; Reader, P. D.

    1972-01-01

    The current status of research and development programs on electrostatic thrusters is reviewed. Current programs that utilize mercury electron-bombardment thrusters range from 5- to 30-cm in diameter. Recent progress on the 5-cm thruster has emphasized durability, with accelerator time exceeding 6300 hours and total time on the rest of the thruster exceeding 8300 hours. Recent progress on the 30-cm thruster has been outstanding in dished-grid accelerator systems. Ion beams up to 5 amperes have been obtained for short periods with 1000 volts net accelerating potential difference. The cesium electron-bombardment and cesium contact programs are also described.

  15. Theta-Pinch Thruster for Piloted Deep Space Exploration

    NASA Technical Reports Server (NTRS)

    LaPointe, Mike R.; Reddy, Dhanireddy (Technical Monitor)

    2000-01-01

    A new high-power propulsion concept that combines a rapidly pulsed theta-pinch discharge with upstream particle reflection by a magnetic mirror was evaluated under a Phase 1 grant awarded through the NASA Institute for Advanced Concepts. Analytic and numerical models were developed to predict the performance of a theta-pinch thruster operated over a wide range of initial gas pressures and discharge periods. The models indicate that a 1 m radius, 10 m long thruster operated with hydrogen propellant could provide impulse-bits ranging from 1 N-s to 330 N-s with specific impulse values of 7,500 s to 2,500 s, respectively. A pulsed magnetic field strength of 2 T is required to compress and heat the preionized hydrogen over a 10(exp -3) second discharge period, with about 60% of the heated plasma exiting the chamber each period to produce thrust. The unoptimized thruster efficiency is low, peaking at approximately 16% for an initial hydrogen chamber pressure of 100 Torr. The specific impulse and impulse-bit at this operating condition are 3,500 s and 90 N-s, respectively, and the required discharge energy is approximately 9x10(exp 6) J. For a pulse repetition rate of 10 Hz, the engine would produce an average thrust of 900 N at 3,500 s specific impulse. Combined with the electrodeless nature of the device, these performance parameters indicate that theta-pinch thrusters could provide unique, long-life propulsion systems for piloted deep space mission applications.

  16. 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.

  17. 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

  18. Engineering Risk Assessment of Space Thruster Challenge Problem

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  19. 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.

  20. Development of a multiplexed electrospray micro-thruster with post-acceleration and beam containment

    NASA Astrophysics Data System (ADS)

    Lenguito, G.; Gomez, A.

    2013-10-01

    We report the development of a compact thruster based on Multiplexed ElectroSprays (MES). It relied on a microfabricated Si array of emitters coupled with an extractor electrode and an accelerator electrode. The accelerator stage was introduced for two purposes: containing beam opening and avoiding electrode erosion due to droplet impingement, as well as boosting specific impulse and thrust. Multiplexing is generally necessary as a thrust multiplier to reach eventually the level required (O(102) μN) by small satellites. To facilitate system optimization and debugging, we focused on a 7-nozzle MES device and compared its performance to that of a single emitter. To ensure uniformity of operation of all nozzles their hydraulic impedance was augmented by packing them with micrometer-size beads. Two propellants were tested: a solution of 21.5% methyl ammonium formate in formamide and the better performing pure ionic liquid ethyl ammonium nitrate (EAN). The 7-MES device spraying EAN at ΔV = 5.93 kV covered a specific impulse range from 620 s to 1900 s and a thrust range from 0.6 μN to 5.4 μN, at 62% efficiency. Remarkably, less than 1% of the beam was demonstrated to impact on the accelerator electrode, which bodes well for long-term applications in space.

  1. MOA - The Magnetic Field Amplified Thruster, a Novel Concept for a Pulsed Plasma Accelerator

    SciTech Connect

    Frischauf, Norbert; Hettmer, Manfred; Grassauer, Andreas; Bartusch, Tobias; Koudelka, Otto

    2008-01-21

    More than 60 years after the later Nobel laureate Hannes Alfven had published a letter stating that oscillating magnetic fields can accelerate ionised matter via magneto-hydrodynamic interactions in a wave like fashion, the technical implementation of Alfven waves for propulsive purposes has been proposed, patented and examined for the first time by a group of inventors. The name of the concept is MOA - Magnetic field Oscillating Amplified thruster. Based on computer simulations, MOA is a highly flexible propulsion system, whose performance parameters might easily be adapted, by changing the mass flow and/or the power level. As such the system is capable to deliver a maximum specific impulse of 13116 s (12.87 mN) at a power level of 11.16 kW, using Xe as propellant, but can also be attuned to provide a thrust of 236.5 mN (2411 s) at 6.15 kW of power. While space propulsion is expected to be the prime application for MOA and is supported by numerous applications such as Solar and/or Nuclear Electric Propulsion or even as an 'afterburner system' for Nuclear Thermal Propulsion, other terrestrial applications can be thought of as well, making the system highly suited for a common space-terrestrial application research and utilisation strategy. This paper presents the recent developments of the MOA Thruster R and D activities at QASAR (www.qasar.at), the company in Vienna, which has been set up to further develop and test the Alfven wave technology and its applications.

  2. On the capabilities of nano electrokinetic thrusters for space propulsion

    NASA Astrophysics Data System (ADS)

    Diez, F. J.; Hernaiz, G.; Miranda, J. J.; Sureda, M.

    2013-02-01

    A theoretical analysis considering the capabilities of nano electrokinetic thrusters for space propulsion is presented. The work describes an electro-hydro-dynamic model of the electrokinetic flow in nano-channels and represents the first attempt to exploit the advantages of the electrokinetic effect as the basis for a new class of nano-scale thrusters suitable for space propulsion. Among such advantages are their small volume, fundamental simplicity, overall low mass, and actuation efficiency. Their electrokinetic efficiency is affected by the slip length, surface charge, pH and molarity. These design variables are analyzed and optimized for the highest electrokinetic performance inside nano-channels. The optimization is done for power consumption, thrust and specific impulse resulting in high theoretical efficiency ˜99% with corresponding high thrust-to-power ratios. Performance curves are obtained for the electrokinetic design variables showing that high molarity electrolytes lead to high thrust and specific impulse values, whereas low molarities provide highest thrust-to-power ratios and efficiencies. A theoretically designed 100 nm wide by 1 μm long emitter optimized using the ideal performance charts developed would deliver thrusts from 5 to 43 μN, specific impulse from 60 to 210 s, and would have power consumption between 1-15 mW. It should be noted that although this is a detail analytical analysis no prototypes exist and any future experimental work will face challenges that could affect the final performance. By designing an array composed of thousands of these single electrokinetic emitters, it would result in a flexible and scalable propulsion system capable of providing a wide range of thrust control for different mission scenarios and maintaining very high efficiencies and thrust-to-power ratio by varying the number of emitters in use at any one time.

  3. Optimum performance of MHD-augumented chemical rocket thrusters for space propulsion applications

    SciTech Connect

    Schulz, R.J.; Chapman, J.N.

    1995-12-31

    The use of magnetohydrodynamic (MHD) acceleration of a chemical rocket exhaust stream, to augment the thrust of small, space-propulsion type chemical thrusters was examined, with the purpose of identifying {open_quotes}optimum{close_quotes} performance. Optimum performance is defined herein as the highest spacecraft acceleration levels with concurrent highest specific impulse, that the hybrid propulsion system can generate, given a fixed mass flow of propellant and fixed chamber pressure (150 psia). The exhaust nozzle-MHD channel selected was of the simplest kind, a three-segmented Faraday generator, for simplicity in design, manufacture, and power control circuit assembly. The channel expanded in only one plane or direction, the plane intersecting the electrodes. The distance between the side walls was fixed. Three different fuel oxidizer combinations were investigated: H{sub 2} - O{sub 2}, fuel oil - O{sub 2}, and hydrazine - nitrogen tetroxide. These represent the spectrum of typical liquid rocket propellants. The fraction of the propellant flow representing potassium, as K{sub 2}CO{sub 3}, was kept constant at 1/2 percent of the total propellant flow. The results of the study verify that the MHD-augmented chemical thruster will be an important propulsion system option for space missions requiring accelerations of the order of milli-gravities with specific impulses of the order of 4,000 seconds. The system study showed that a 3-segmented, diverging Faraday channel with about a 2{degrees} divergence angle, enclosed by a 4 Tesla magnet, was capable of providing exhaust gas exit velocities of the order of 40000 m/s for all three propellant combinations. Hence, a hybrid propulsion system of the type identified here is capable of providing thrusts of the order of 400 Newtons, spacecraft accelerations of the order 2 milli-gravities, with electric power requirements of about 2.4 megawatts, based on propellant total mass flow rates of about 10 grams per second.

  4. 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.

  5. A high power ion thruster for deep space missions.

    PubMed

    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.

  6. Systems and methods for cylindrical hall thrusters with independently controllable ionization and acceleration stages

    DOEpatents

    Diamant, Kevin David; Raitses, Yevgeny; Fisch, Nathaniel Joseph

    2014-05-13

    Systems and methods may be provided for cylindrical Hall thrusters with independently controllable ionization and acceleration stages. The systems and methods may include a cylindrical channel having a center axial direction, a gas inlet for directing ionizable gas to an ionization section of the cylindrical channel, an ionization device that ionizes at least a portion of the ionizable gas within the ionization section to generate ionized gas, and an acceleration device distinct from the ionization device. The acceleration device may provide an axial electric field for an acceleration section of the cylindrical channel to accelerate the ionized gas through the acceleration section, where the axial electric field has an axial direction in relation to the center axial direction. The ionization section and the acceleration section of the cylindrical channel may be substantially non-overlapping.

  7. Space Shuttle reaction control system thruster metal nitrate removal and characterization

    NASA Technical Reports Server (NTRS)

    Saulsberry, R. L.; Mccartney, P. A.

    1993-01-01

    The Space Shuttle hypergolic primary reaction control system (PRCS) thrusters continue to fail-leak or fail-off at a rate of approximately 1.5 per flight, attributed primarily to metal nitrate formation in the nitrogen tetroxide (N2O4) pilot operated valves (POV's). The failures have continued despite ground support equipment (GSE) and subsystem operational improvements. As a result, the Johnson Space Center (JSC) White Sands Test Facility (WSTF) performed a study to characterize the contamination in the N204 valves. This study prompted the development and implementation of a highly successful flushing technique using deionized (DI) water and gaseous nitrogen (GN2) to remove the contamination while minimizing Teflon seat damage. Following flushing a comprehensive acceptance test is performed before the thruster is deemed recovered. Between the time WSTF was certified to process flight thrusters (March 1992) and September 1993, a 68 percent thruster recovery rate was achieved. The contamination flushed from these thrusters was analyzed and has provided insight into the corrosion process, which is reported in this publication. Additionally, the long-term performance of 24 flushed thrusters installed in the WSTF Fleet Leader Shuttle reaction control subsystem (RCS) test articles is being assessed. WSTF continues to flush flight and test article thrusters and compile data to investigate metal nitrate formation characteristics in leaking and nonleaking valves.

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

    SciTech Connect

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

    2005-03-16

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

  9. 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.

  10. Magnetic Shielding of the Acceleration Channel Walls in a Long-Life Hall Thruster

    NASA Technical Reports Server (NTRS)

    Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.; Goebel, Dan M.; de Grys, Kristi; Mathers, Alex

    2010-01-01

    In a Qualification Life Test (QLT) of the BPT-4000 Hall thruster that recently accumulated greater than 10,000 h it was found that the erosion of the acceleration channel practically stopped after approximately 5,600 h. Numerical simulations of this thruster using a 2-D axisymmetric, magnetic field-aligned-mesh (MFAM) plasma solver reveal that the process that led to this significant reduction of the erosion was multifaceted. It is found that when the channel receded from its early-in-life geometry to its steady-state configuration several changes in the near-wall plasma and sheath were induced by the magnetic field that, collectively, constituted an effective shielding of the walls from any significant ion bombardment. Because all such changes in the behavior of the ionized gas near the eroding surfaces were caused by the topology of the magnetic field there, we term this process "magnetic shielding."

  11. The Development of Plasma Thrusters and Its Importance for Space Technology and Science Education at University of Brasilia

    NASA Astrophysics Data System (ADS)

    Ferreira, Jose Leonardo; Calvoso, Lui; Gessini, Paolo; Ferreira, Ivan

    Since 2004 The Plasma Physics Laboratory of University of Brasilia (Brazil) is developing Hall Plasma Thurusters for Satellite station keeping and orbit control. The project is supported by CNPq, CAPES, FAP DF and from The Brazillian Space Agency-AEB. The project is part of The UNIESPAÇO Program for Space Activities Development in Brazillian Universities. In this work we are going to present the highlights of this project together with its vital contribution to include University of Brasilia in the Brazillian Space Program. Electric propulsion has already shown, over the years, its great advantages in being used as main and secondary thruster system of several space mission types. Between the many thruster concepts, one that has more tradition in flying real spacecraft is the Hall Effect Thruster (HET). These thrusters, first developed by the USSR in the 1960s, uses, in the traditional design, the radial magnetic field and axial electric field to trap electrons, ionize the gas and accelerate the plasma to therefore generate thrust. In contrast to the usual solution of using electromagnets to generate the magnetic field, the research group of the Plasma Physics Laboratory of University of Brasília has been working to develop new models of HETs that uses combined permanent magnets to generate the necessary magnetic field, with the main objective of saving electric power in the final system design. Since the beginning of this research line it was developed and implemented two prototypes of the Permanent Magnet Hall Thruster (PMHT). The first prototype, called P-HALL1, was successfully tested with the using of many diagnostics instruments, including, RF probe, Langmuir probe, Ion collector and Ion energy analyzer. The second prototype, P-HALL2, is currently under testing, and it’s planned the increasing of the plasma diagnostics and technology analysis, with the inclusion of a thrust balance, mass spectroscopy and Doppler broadening. We are also developing an

  12. Plasma Diagnostics of a Forward Laser Plasma Accelerated Thruster

    SciTech Connect

    Izumi, Masaya; Horisawa, Hideyuki; Takeda, Akihito; Kimura, Itsuro

    2006-05-02

    Fundamental investigations on plasma diagnostics of a forward laser plasma acceleration employing laser-foil interactions were conducted for an Al-foil target irradiated with an Nd:YAG laser of 1J/pulse with pulse-width of 10nsec. A time-of-flight measurement was also conducted to evaluate ion speeds. In addition, temporal evolutions of electron temperatures and densities were evaluated with electrostatic probes and spectroscopic diagnostics. Moreover, a preliminary one-dimensional particle-in-cell (PIC) simulation was conducted to elucidate acceleration mechanisms. From the results, it was shown that a speed of ions in a forward direction were about 135 km/sec, respectively. Also it was shown that the plasma temperature and density were about 2.5{approx}3 eV and 1010 cm-3.

  13. Ion Thruster and Power Processor Developed for the Deep Space 1 Mission

    NASA Technical Reports Server (NTRS)

    Sovey, James S.

    1999-01-01

    The NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) Program has provided a single-string primary propulsion system to NASA's Deep Space 1 spacecraft. This spacecraft will carry about 81 kg of xenon propellant for the ion thruster, which can be throttled down from 2.3 to 0.5 kW as the spacecraft moves away from the Sun. The propellant load will provide about 20 months of propulsion at the one-half power throttle setpoint of 1.2 kW. This mission will validate the 2.5-kW ion propulsion system and will fly by the asteroid 1992 KD in 1999. If funding permits, Deep Space 1 also will encounter comets Wilson-Harrington and Borrelly in 2001. NASA Lewis Research Center's On-Board Propulsion Branch was responsible for the development of the 30-cm-diameter ion thruster, the 2.5-kW power processor unit (PPU), and the Digital Control and Interface Unit (DCIU) that controls the PPU/thruster/feed system and provides data and recovery from fault conditions. Lewis transferred the thruster and PPU technologies to the Hughes Electron Dynamics Division, which was selected to build two sets of flight thrusters, as well as the PPU's and DCIU's. Hughes subcontracted the DCIU development to Spectrum Astro Incorporated. The Jet Propulsion Laboratory (JPL) was primarily responsible for the NSTAR project management, thruster lifetests, the feed system, diagnostics, and the propulsion subsystem integration. A total of four engineering model thrusters and three breadboard PPU's were built, integrated, and tested. More than 50 development tests were conducted along with thruster design verification tests of 2000 and 1000 hours. In addition, an 8000-hr life demonstration test was successfully completed and demonstrated wear-rates consistent with full-power lifetimes in excess of 12,000 hours.

  14. Hydrogen-oxygen space shuttle ACPS thruster technology review

    NASA Technical Reports Server (NTRS)

    Gregory, J. W.; Herr, P. N.

    1972-01-01

    The generation of technology for injectors, cooled thrust chambers, valves, and ignition systems is discussed. The thrusters are designed to meet a unique and stringent set of requirements, including: long life for 100 mission reuses, high performance, light weight, ability to provide long duration firings as well as small impulse bits, ability to operate over wide ranges of propellant inlet conditions and to withstand reentry heating. The program has included evaluation of thrusters designed for ambient temperature and cold gaseous propellants at the vehicle interface.

  15. 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.

  16. Liquid-metal-fed Pulsed Plasma Thrusters for In-space Propulsion

    NASA Technical Reports Server (NTRS)

    Markusic, Thomas E.

    2004-01-01

    Liquid metal propellants may provide a path toward more reliable and efficient pulsed plasma thrusters (PPTs). Conceptual thruster designs which eliminate the need for high current switches and propellant metering valves are described. Propellant loading techniques are suggested that show promise to increase thruster propellant utilization, dynamic, and electrical efficiency. Calibration results from a compact, electromagnetically-pumped propellant feed system are presented. Results for lithium and gallium propellants show capability to meter propellant at flow rates up to 10 +/- 0.1 mg/s. Experiments investigating the initiation of arc discharges using liquid metal droplets are presented. High speed photography and laser interferometry provide spatially and temporally resolved information on the decomposition of liquid metal droplets , and the evolution of the accelerating current channel.

  17. Performance of Solar Electric Powered Deep Space Missions Using Hall Thruster Propulsion

    NASA Technical Reports Server (NTRS)

    Witzberger, Kevin E.; Manzella, David

    2006-01-01

    Power limited, low-thrust trajectories were assessed for missions to Jupiter, Saturn, and Neptune utilizing a single Venus Gravity Assist (VGA) and a primary propulsion system based on either a 3-kW high voltage Hall thruster, of the type being developed by the NASA In-Space Propulsion Technology Program, or an 8-kW variant of this thruster. These Hall thrusters operate with specific impulses below 3,000 seconds. A trade study was conducted to examine mission parameters that include: net delivered mass (NDM), beginning-of-life (BOL) solar array power, heliocentric transfer time, required launch vehicle, number of operating thrusters, and throttle profile. The top performing spacecraft configuration was defined to be the one that delivered the highest mass for a range of transfer times. In order to evaluate the potential future benefit of using next generation Hall thrusters as the primary propulsion system, comparisons were made with the advanced state-of-the-art (ASOA), 7-kW, 4,100 second NASA's Evolutionary Xenon Thruster (NEXT) for the same mission scenarios. For the BOL array powers considered in this study (less than 30 kW), the results show that the performance of the Hall thrusters, relative to NEXT, is largely dependant on the performance capability of the launch vehicle, and that at least a 10 percent performance gain, equating to at least an additional 200 kg dry mass at each target planet, is achieved over the higher specific impulse NEXT when launched on an Atlas 551.

  18. Accelerated testing of space batteries

    NASA Technical Reports Server (NTRS)

    Mccallum, J.; Thomas, R. E.; Waite, J. H.

    1973-01-01

    An accelerated life test program for space batteries is presented that fully satisfies empirical, statistical, and physical criteria for validity. The program includes thermal and other nonmechanical stress analyses as well as mechanical stress, strain, and rate of strain measurements.

  19. Expanding the Capabilities of the Pulsed Plasma Thruster for In-Space and Atmospheric Operation

    NASA Astrophysics Data System (ADS)

    Johnson, Ian Kronheim

    Of all in-space propulsion systems to date, the Pulsed Plasma Thruster (PPT) is unique in its simplicity and wide range of operational parameters. This study examined multiple uses of the thruster for in-space and atmospheric propulsion, as well as the creation of a CubeSat satellite and atmospheric airship as test beds for the thruster. The PPT was tested as a solid-propellant feed source for the High Power Helicon Thruster, a compact plasma source capable of generating order of magnitude higher plasma densities than comparable power level systems. Replacing the gaseous feed system reduced the thruster size and complexity, as well as allowing for extremely discrete discharges, minimizing the influence of wall effects. Teflon (C2F4) has been the traditional propellant for PPTs due to a high exhaust velocity and ability to ablate without surface modification over long durations. A number of alternative propellants, including minerals and metallics commonly found on asteroids, were tested for use with the PPT. Compounds with significant fractions of sulfur showed the highest performance increase, with specific thrusts double that of Teflon. A PPT with sulfur propellant designed for CubeSat operation, as well as the subsystems necessary for autonomous operation, was built and tested in the laboratory. The PPT was modified for use at atmospheric pressures where the impulse was well defined as a function of the discharge chamber volume, capacitor energy, and background pressure. To demonstrate that the air-breathing PPT was a viable concept the device was launched on two atmospheric balloon flights.

  20. 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

  1. Deflagration plasma thruster

    NASA Technical Reports Server (NTRS)

    Cheng, D. Y.; Chang, C. N.

    1984-01-01

    This paper introduces the application of the magnetized plasma deflagration process to space propulsion. The deflagration process has the unique capability of efficiently converting input energy into kinetic energy in the accelerating direction. To illustrate the totally divergent characters of 'snowplow' detonation and deflagration discharges, examples of the differences between deflagration and detonation 'snowplow' discharges are expressed in terms of current densities, temperature, and particle velocities. Magnetic field profiles of the deflagration mode of discharges are measured. Typical attainable plasma characteristics are described in terms of velocity, electron temperature, and density, as well as measurement techniques. Specific impulses measured by piezo-electric probe and pendulum methods are presented. The influence of the transmission line in the discharge circuits on plasma velocity is measured by means of a microwave time-of-flight method. The results for the deflagration thruster are compared with other space thrusters. Further research areas are identified.

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

    NASA Astrophysics Data System (ADS)

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

    2016-07-01

    The Plasma Physics Laboratory of UnB has been developing a Permanent Magnet Hall Thruster (PHALL) for the UNIESPAÇO program, part of the Space Activities Program conducted by AEB- The Brazillian Space Agency since 2004. Electric propulsion is now a very successful method for primary and secondary propulsion systems. It is essential for several existing geostationary satellite station keeping systems and for deep space long duration solar system missions, where the thrusting system can be designed to be used on orbit transfer maneuvering and/or for satellite attitude control in long term space missions. Applications of compact versions of Permanent Magnet Hall Thrusters on future brazillian space missions are needed and foreseen for the coming years beginning with the use of small divergent cusp field (DCFH) Hall Thrusters type on CUBESATS ( 5-10 kg , 1W-5 W power consumption) and on Micro satellites ( 50- 100 kg, 10W-100W). Brazillian (AEB) and German (DLR) space agencies and research institutions are developing a new rocket dedicated to small satellite launching. The VLM- Microsatellite Launch Vehicle. The development of PHALL compact versions can also be important for the recently proposed SBG system, a future brazillian geostationary satellite system that is already been developed by an international consortium of brazillian and foreign space industries. The exploration of small bodies in the Solar System with spacecraft has been done by several countries with increasing frequency in these past twenty five years. Since their historical beginning on the sixties, most of the Solar System missions were based on gravity assisted trajectories very much depended on planet orbit positioning relative to the Sun and the Earth. The consequence was always the narrowing of the mission launch window. Today, the need for Solar System icy bodies in situ exploration requires less dependence on gravity assisted maneuvering and new high precision low thrust navigation methods

  3. A study of cylindrical Hall thruster for low power space applications

    SciTech Connect

    Y. Raitses; N.J. Fisch; K.M. Ertmer; C.A. Burlingame

    2000-07-27

    A 9 cm cylindrical thruster with a ceramic channel exhibited performance comparable to the state-of-the-art Hall thrusters at low and moderate power levels. Significantly, its operation is not accompanied by large amplitude discharge low frequency oscillations. Preliminary experiments on a 2 cm cylindrical thruster suggest the possibility of a high performance micro Hall thruster.

  4. High-Power Magnetoplasmadynamic Thruster Being Developed

    NASA Technical Reports Server (NTRS)

    LaPointe, Michael R.

    2001-01-01

    High-power electromagnetic thrusters have been proposed as primary in-space propulsion options for several of the bold new interplanetary and deep space missions envisioned by the Human Exploration and Development of Space (HEDS) Strategic Enterprise. As the lead center for electric propulsion, the NASA Glenn Research Center is actively involved in the design, development, and testing of high-power electromagnetic technologies to meet these demanding mission requirements. One concept of particular interest is the magnetoplasmadynamic (MPD) thruster, shown schematically in the preceding figure. In its basic form, the MPD thruster consists of a central cathode surrounded by a concentric cylindrical anode. A high-current arc is struck between the anode and cathode, which ionizes and accelerates a gas (plasma) propellant. In the self-field version of the thruster, an azimuthal magnetic field generated by the current returning through the cathode interacts with the radial discharge current flowing through the plasma to produce an axial electromagnetic body force, providing thrust. In applied field-versions of the thruster, a magnetic field coil surrounding the anode is used to provide additional radial and axial magnetic fields that can help stabilize and accelerate the plasma propellant. The following figure shows an experimental megawatt-class MPD thruster developed at Glenn. The MPD thruster is fitted inside a magnetic field coil, which in turn is mounted on a thrust stand supported by thin metal flexures. A calibrated position transducer is used to determine the force provided by the thruster as a function of thrust stand displacement. Power to the thruster is supplied by a 250-kJ capacitor bank, which provides up to 30- MW to the thruster for a period of 2 msec. This short period of time is sufficient to establish thruster performance similar to steady-state operation, and it allows a number of thruster designs to be quickly and economically evaluated. In concert

  5. 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

  6. Accelerated testing of space mechanisms

    NASA Technical Reports Server (NTRS)

    Murray, S. Frank; Heshmat, Hooshang

    1995-01-01

    This report contains a review of various existing life prediction techniques used for a wide range of space mechanisms. Life prediction techniques utilized in other non-space fields such as turbine engine design are also reviewed for applicability to many space mechanism issues. The development of new concepts on how various tribological processes are involved in the life of the complex mechanisms used for space applications are examined. A 'roadmap' for the complete implementation of a tribological prediction approach for complex mechanical systems including standard procedures for test planning, analytical models for life prediction and experimental verification of the life prediction and accelerated testing techniques are discussed. A plan is presented to demonstrate a method for predicting the life and/or performance of a selected space mechanism mechanical component.

  7. Laser Fine-Adjustment Thruster For Space Vehicles

    SciTech Connect

    Rezunkov, Yu. A.; Egorov, M. S.; Repina, E. V.; Safronov, A. L.; Rebrov, S. G.

    2010-05-06

    To the present time, a few laser propulsion engine devices have been developed by using dominant mechanisms of laser propulsion. Generally these mechanisms are laser ablation, laser breakdown of gases, and laser detonation waves that are induced due to extraction of the internal energy of polymer propellants. In the paper, we consider the Aero-Space Laser Propulsion Engine (ASLPE) developed earlier, in which all of these mechanisms are realized via interaction of laser radiation with polymers both in continuous wave (CW) and in repetitively pulsed modes of laser operation. The ASLPE is considered to be exploited as a unit of a laser propulsion device being arranged onboard space vehicles moving around the Earth or in interplanetary missions and intended to correct the vehicles orbits. To produce a thrust, a power of the solar pumped lasers designed to the present time is considered in the paper. The problem of increasing the efficiency of the laser propulsion device is analyzed as applied to space missions of vehicles by optimizing the laser propulsion propellant composition.

  8. Hydrogen-oxygen auxiliary propulsion for the space shuttle. Volume 2: Low pressure thrusters

    NASA Technical Reports Server (NTRS)

    1973-01-01

    An abbreviated program was conducted to investigate igniter, injector, and thrust chamber technology for a 10.3 N/cm2 (15 psia) chamber pressure, 6660 N (1500 lbf) gaseous H2/O2 APS thruster for the Space Shuttle Vehicle. Successful catalytic igniter tests were conducted with ambient and cold propellants. Injector testing with a heat sink chamber (MR = 2.5, area ratio = 5.0) gave a measured specific impulse of 386 sec with 11% of the fuel used as film coolant. This coolant flow rate was demonstrated to be more than adequate to cool a spun adiabatic wall, flightweight thrust chamber.

  9. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2016-07-01

    The Plasma Physics Laboratory of UnB has been developing a Permanent Magnet Hall Thruster (PHALL) for the UNIESPAÇO program, part of the Space Activities Program conducted by AEB- The Brazillian Space Agency since 2004. Electric propulsion is now a very successful method for primary and secondary propulsion systems. It is essential for several existing geostationary satellite station keeping systems and for deep space long duration solar system missions, where the thrusting system can be designed to be used on orbit transfer maneuvering and/or for satellite attitude control in long term space missions. Applications of compact versions of Permanent Magnet Hall Thrusters on future brazillian space missions are needed and foreseen for the coming years beginning with the use of small divergent cusp field (DCFH) Hall Thrusters type on CUBESATS ( 5-10 kg , 1W-5 W power consumption) and on Micro satellites ( 50- 100 kg, 10W-100W). Brazillian (AEB) and German (DLR) space agencies and research institutions are developing a new rocket dedicated to small satellite launching. The VLM- Microsatellite Launch Vehicle. The development of PHALL compact versions can also be important for the recently proposed SBG system, a future brazillian geostationary satellite system that is already been developed by an international consortium of brazillian and foreign space industries. The exploration of small bodies in the Solar System with spacecraft has been done by several countries with increasing frequency in these past twenty five years. Since their historical beginning on the sixties, most of the Solar System missions were based on gravity assisted trajectories very much depended on planet orbit positioning relative to the Sun and the Earth. The consequence was always the narrowing of the mission launch window. Today, the need for Solar System icy bodies in situ exploration requires less dependence on gravity assisted maneuvering and new high precision low thrust navigation methods

  11. Space simulation experiments on reaction control system thruster plumes.

    NASA Technical Reports Server (NTRS)

    Cassidy, J. F.

    1972-01-01

    A space simulation procedure was developed for studying rocket plume contamination effects using a 5-lb bipropellant reaction control system thrustor. Vacuum chamber pressures of 0.00003 torr (70 miles altitude) were achieved with the thrustor firing in pulse trains consisting of eight pulses - 50 msec on, 100 msec off, and seven minutes between pulse trains. The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid-helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.

  12. Conical solar absorber/thruster for space propulsion

    SciTech Connect

    Strumpf, H.J.; Borghese, J.B.; Keating, R.F.

    1995-11-01

    Solar-powered space propulsion uses solar heating of a propellant such as hydrogen to impart thrust to a rocket when the hydrogen exists through an appropriately designed nozzle. Because of the low molecular weight of hydrogen, exhaust velocities, and hence specific impulses, can potentially be much greater than for chemical combustion of fuel. A very efficient solar thermal absorber design has been developed. The design consists of two interwound helical coils of rhenium tubing, through which the propellant flows to be heated before being exhausted out a rhenium nozzle. The conical absorbing surface is configured to conform to the extreme solar rays from a solar concentrator; i.e., the receiver apex angle is designed to match the concentrator apex angle. This shape helps to minimize the amount of reflected or emitted energy lost through the receiver aperture.

  13. Laser-Supported Detonation Concept as a Space Thruster

    SciTech Connect

    Fujiwara, Toshi; Miyasaka, Takeshi

    2004-03-30

    Similar to the concept of pulse detonation engine (PDE), a detonation generated in the 'combustion chamber' due to incoming laser absorption can produce the thrust basically much higher than the one that a laser-supported deflagration wave can provide. Such a laser-supported detonation wave concept has been theoretically studied by the first author for about 20 years in view of its application to space propulsion. The entire work is reviewed in the present paper. The initial condition for laser absorption can be provided by increasing the electron density using electric discharge. Thereafter, once a standing/running detonation wave is formed, the laser absorption can continuously be performed by the classical absorption mechanism called Inverse Bremsstrahlung behind a strong shock wave.

  14. 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.

  15. Permanent magnet Hall Thrusters development and applications on future brazilian space missions

    NASA Astrophysics Data System (ADS)

    Ferreira, Jose Leonardo; Martins, Alexandre A.; Miranda, Rodrigo; Schelling, Adriane; de Souza Alves, Lais; Gonçalves Costa, Ernesto; de Oliveira Coelho Junior, Helbert; Castelo Branco, Artur; de Oliveira Lopes, Felipe Nathan

    2015-10-01

    The Plasma Physics Laboratory (PPLUnB) has been developing a Permanent Magnet Hall Thruster (PHALL) for the Space Research Program for Universities (UNIESPAÇO), part of the Brazilian Space Activities Program (PNAE) since 2004. The PHALL project consists on a plasma source design, construction and characterization of the Hall type that will function as a plasma propulsion engine and characterized by several plasma diagnostics sensors. PHALL is based on a plasma source in which a Hall current is generated inside a cylindrical annular channel with an axial electric field produced by a ring anode and a radial magnetic field produced by permanent magnets. In this work it is shown a brief description of the plasma propulsion engine, its diagnostics instrumentation and possible applications of PHALL on orbit transfer maneuvering for future Brazilian geostationary satellite space missions.

  16. LLNL contributions to MPD thrusters for SEI

    NASA Technical Reports Server (NTRS)

    Hooper, Edwin Bickford

    1991-01-01

    Some of the topics covered with respect to the Lawrence Livermore National Laboratory's (LLNL's) contributions to Magnetoplasmadynamic (MPD) Thrusters for the Space Exploration Initiative (SEI) include: an IR camera, plasma-induced erosion/redeposition, the Mirror Fusion Test Facility-B (MFTF-B), the Thruster Lifetime Test Facility, the RACE Compact Torus Accelerator Facility, and a RACE program summary. Some of the other topics addressed include: flux contours for HAM simulation, comparison of RACE data of plasma ring formation with the HAM 2-D magnetohydrodynamic code, and the 2-D Ring Acceleration Code (TRAC).

  17. Metallographic Preparation of Space Shuttle Reaction Control System Thruster Electron Beam Welds for Electron Backscatter Diffraction

    NASA Technical Reports Server (NTRS)

    Martinez, James

    2011-01-01

    A Space Shuttle Reaction Control System (RCS) thruster failed during a firing test at the NASA White Sands Test Facility (WSTF), Las Cruces, New Mexico. The firing test was being conducted to investigate a previous electrical malfunction. A number of cracks were found associated with the fuel closure plate/injector assembly (Fig 1). The firing test failure generated a flight constraint to the launch of STS-133. A team comprised of several NASA centers and other research institutes was assembled to investigate and determine the root cause of the failure. The JSC Materials Evaluation Laboratory was asked to compare and characterize the outboard circumferential electron beam (EB) weld between the fuel closure plate (Titanium 6Al-4V) and the injector (Niobium C-103 alloy) of four different RCS thrusters, including the failed RCS thruster. Several metallographic challenges in grinding/polishing, and particularly in etching were encountered because of the differences in hardness, ductility, and chemical resistance between the two alloys and the bimetallic weld. Segments from each thruster were sectioned from the outboard weld. The segments were hot-compression mounted using a conductive, carbon-filled epoxy. A grinding/polishing procedure for titanium alloys was used [1]. This procedure worked well on the titanium; but a thin, disturbed layer was visible on the niobium surface by means of polarized light. Once polished, each sample was micrographed using bright field, differential interference contrast optical microscopy, and scanning electron microscopy (SEM) using a backscatter electron (BSE) detector. No typical weld anomalies were observed in any of the cross sections. However, areas of large atomic contrast were clearly visible in the weld nugget, particularly along fusion line interfaces between the titanium and the niobium. This prompted the need to better understand the chemistry and microstructure of the weld (Fig 2). Energy Dispersive X-Ray Spectroscopy (EDS

  18. 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.

  19. 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. __________________________________________________

  20. 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.

  1. The NASA Evolutionary Xenon Thruster (NEXT): NASA's Next Step for U.S. Deep Space Propulsion

    NASA Technical Reports Server (NTRS)

    Schmidt, George R.; Patterson, Michael J.; Benson, Scott W.

    2008-01-01

    NASA s Evolutionary Xenon Thruster (NEXT) project is developing next generation ion propulsion technologies to enhance the performance and lower the costs of future NASA space science missions. This is being accomplished by producing Engineering Model (EM) and Prototype Model (PM) components, validating these via qualification-level and integrated system testing, and preparing the transition of NEXT technologies to flight system development. The project is currently completing one of the final milestones of the effort, that is operation of an integrated NEXT Ion Propulsion System (IPS) in a simulated space environment. This test will advance the NEXT system to a NASA Technology Readiness Level (TRL) of 6 (i.e., operation of a prototypical system in a representative environment), and will confirm its readiness for flight. Besides its promise for upcoming NASA science missions, NEXT may have excellent potential for future commercial and international spacecraft applications.

  2. Space experiments with particle accelerators

    SciTech Connect

    Obayashi, T.

    1981-11-01

    The purpose of space experiments with particle accelerators (SEPAC) is to carry out active and interactive experiments on and in the Earth's ionosphere and magnetosphere. It is also intended to make an initial performance test for an overall program of Spacelab/SEPAC experiments. The instruments to be used are an electron beam accelerator, magnetoplasma dynamic arcjet, and associated diagnostic equipment. The accelerators are installed on the pallet, with monitoring and diagnostic observations being made by the gas plume release, beam-monitor TV, and particle-wave measuring instruments also mounted on the pallet. Command and display systems are installed in the module. Three major classes of investigations to be performed are vehicle charge neutralization, beam plasma physics, and beam atmosphere interactions. The first two are mainly onboard plasma physics experiments to measure the effect of phenomena in the vicinity of Spacelab. The last one is concerned with atmospheric modification and is supported by other Spacelab 1 investigations as well as by ground-based, remote sensing observations.

  3. Development of a high power microwave thruster, with a magnetic nozzle, for space applications

    NASA Technical Reports Server (NTRS)

    Power, John L.; Chapman, Randall A.

    1989-01-01

    This paper describes the current development of a high-power microwave electrothermal thruster (MET) concept at the NASA Lewis Research Center. Such a thruster would be employed in space for applications such as orbit raining, orbit maneuvering, station change, and possibly trans-lunar or trans-planetary propulsion of spacecraft. The MET concept employs low frequency continuous wave (CW) microwave power to create and continuously pump energy into a flowing propellant gas at relative high pressure via a plasma discharge. The propellant is heated to very high bulk temperatures while passing through the plasma discharge region and then is expanded through a throat-nozzle assembly to produce thrust, as in a conventional rocket engine. Apparatus, which is described, is being assembled at NASA Lewis to test the MET concept to CW power levels of 30 kW at a frequency of 915 MHz. The microwave energy is applied in a resonant cavity applicator and is absorbed by a plasma discharge in the flowing propellant. The ignited plasma acts as a lossy load, and with optimal tuning, energy absorption efficiencies over 95 percent (based on the applied microwave power) are expected. Nitrogen, helium, and hydrogen will be tested as propellants in the MET, at discharge chamber pressures to 10 atm.

  4. 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.

  5. Study of applied magnetic field magnetoplasmadynamic thrusters with particle-in-cell and Monte Carlo collision. II. Investigation of acceleration mechanisms

    SciTech Connect

    Tang Haibin; Cheng Jiao; Liu Chang; York, Thomas M.

    2012-07-15

    The particle-in-cell method previously described in paper (I) has been applied to the investigation of acceleration mechanisms in applied-field magnetoplasmadynamic thrusters. This new approach is an alternative to magnetohydrodynamics models and allows nonlocal dynamic effects of particles and improved transport properties. It was used to model a 100 kW, steady-state, applied-field, argon magnetoplasmadynamic thruster to study the physical acceleration processes with discharge currents of 1000-1500 A, mass flow rates of 0.025-0.1 g/s and applied magnetic field strengths of 0.034-0.102 T. The total thrust calculations were used to verify the theoretical approach by comparison with experimental data. Investigations of the acceleration model offer an underlying understanding of applied-field magnetoplasmadynamic thrusters, including the following conclusions: (1) swirl acceleration mechanism is the dominant contributor to the plasma acceleration, and self-magnetic, Hall, gas-dynamic, and swirl acceleration mechanisms are in an approximate ratio of 1:10:10:100; (2) the Hall acceleration produced mainly by electron swirl is insensitive to the change of externally applied magnetic field and shows only slight increases when the current is raised; (3) self-magnetic acceleration is normally negligible for all cases, while the gas-dynamic acceleration contribution increases with increasing applied magnetic field strength, discharge current, and mass flow rate.

  6. 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.

  7. Development of D+3He Fusion Electric Thrusters and Power Supplies for Space

    NASA Astrophysics Data System (ADS)

    Morse, Thomas M.

    1994-07-01

    Development of D+3He Fusion Electric Thrusters (FET) and Power Supplies (FPS) should occur at a lunar base because of the following: availability of helium-3, a vacuum better than on Earth, low K in shade reachable by radiant cooling, supply of ``high temp'' superconducting ceramic-metals, and a low G environment. The early FET will be much smaller than an Apollo engine, with specific impulse of 10,000-100,000-s. Solar power and low G will aid early development. To counter the effect of low G on humans, centrifuges will be employed for sleeping and resting. Work will be done by telerobotic view control. The FPS will be of comparable size, and will generate power mainly by having replaceable rectennas, resonant to the fusion synchrotron radiation. FPSs are used for house keeping power and initiating superconduction. Spaceships will carry up to ten FETs and two FPSs. In addition to fusion fuel, the FET will inject H or Li low mass propellant into the fusion chamber. Developing an FET would be difficult on Earth. FET spaceships will park between missions in L1, and an FET Bus will fetch humans/supplies from Moon and Earth. Someday FETs, with rocket assist, will lift spaceships from Earth, and make space travel to planets far cheaper, faster, and safer, than at present. Too long a delay due to the space station, or the huge cost of getting into space by current means, will damage the morale of the space program.

  8. 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

  9. 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

  10. 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

  11. 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.

  12. Acceleration Environment of the International Space Station

    NASA Technical Reports Server (NTRS)

    McPherson, Kevin; Kelly, Eric; Keller, Jennifer

    2009-01-01

    Measurement of the microgravity acceleration environment on the International Space Station has been accomplished by two accelerometer systems since 2001. The Microgravity Acceleration Measurement System records the quasi-steady microgravity environment, including the influences of aerodynamic drag, vehicle rotation, and venting effects. Measurement of the vibratory/transient regime, comprised of vehicle, crew, and equipment disturbances, has been accomplished by the Space Acceleration Measurement System-II. Until the arrival of the Columbus Orbital Facility and the Japanese Experiment Module, the location of these sensors, and therefore, the measurement of the microgravity acceleration environment, has been limited to within the United States Laboratory. Japanese Aerospace Exploration Agency has developed a vibratory acceleration measurement system called the Microgravity Measurement Apparatus which will be deployed within the Japanese Experiment Module to make distributed measurements of the Japanese Experiment Module's vibratory acceleration environment. Two Space Acceleration Measurement System sensors from the United States Laboratory will be re-deployed to support vibratory acceleration data measurement within the Columbus Orbital Facility. The additional measurement opportunities resulting from the arrival of these new laboratories allows Principal Investigators with facilities located in these International Space Station research laboratories to obtain microgravity acceleration data in support of their sensitive experiments. The Principal Investigator Microgravity Services project, at NASA Glenn Research Center, in Cleveland, Ohio, has supported acceleration measurement systems and the microgravity scientific community through the processing, characterization, distribution, and archival of the microgravity acceleration data obtained from the International Space Station acceleration measurement systems. This paper summarizes the PIMS capabilities available

  13. 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.

  14. Thruster endurance test

    NASA Technical Reports Server (NTRS)

    Collett, C.

    1976-01-01

    A test system was built and several short term tests were completed. The test system included, in addition to the 30-cm ion thruster, a console for powering the thruster and monitoring performance, a vacuum facility for simulating a space environment, and a storage and feed system for the thruster propellant. This system was used to perform three short term tests (one 100-hour and two 500-hour tests), an 1108-hour endurance test which was aborted by a vacuum facility failure, and finally the 10,000-hour endurance test. In addition to the two 400 series thrusters which were used in the short term and 1100-hour tests, four more 400 series thrusters were fabricated, checked out, and delivered to NASA. Three consoles similar to the one used in the test program were also fabricated and delivered.

  15. NASA Marshall Space Flight Center Tri-gas Thruster Performance Characterization

    NASA Technical Reports Server (NTRS)

    Dorado, Vanessa; Grunder, Zachary; Schaefer, Bryce; Sung, Meagan; Pedersen, Kevin

    2013-01-01

    Historically, spacecraft reaction control systems have primarily utilized cold gas thrusters because of their inherent simplicity and reliability. However, cold gas thrusters typically have a low specific impulse. It has been determined that a higher specific impulse can be achieved by passing a monopropellant fluid mixture through a catalyst bed prior to expulsion through the thruster nozzle. This research analyzes the potential efficiency improvements from using tri-gas, a mixture of hydrogen, oxygen, and an inert gas, which in this case is helium. Passing tri-gas through a catalyst causes the hydrogen and oxygen to react and form water vapor, ultimately heating the exiting fluid and generating a higher specific impulse. The goal of this project was to optimize the thruster performance by characterizing the effects of varying several system components including catalyst types, catalyst lengths, and initial catalyst temperatures.

  16. The interaction of the atmosphere with the space shuttle thruster plume: The NH (A-X) 336-nm emission

    NASA Astrophysics Data System (ADS)

    Viereck, Rodney A.; Murad, Edmond; Knecht, David J.; Pike, Charles P.; Bernstein, Lawrence S.; Elgin, James B.; Broadfoot, A. Lyle

    1996-03-01

    Observations of the optical emissions from the space shuttle's thrusters have been examined. Particular attention has been paid to the interaction of the thruster plume with the atmosphere. Emissions from CN, CH, C2, HNO, and NO2 have been observed near the nozzle of the thruster in the vacuum core region of the plume, but these emissions are the direct result of the combustion process. Other emissions including OI and NH have been observed in the downstream region of the plume, where the plume effluents interact with the atmosphere. The NH emission is one of the most dominant UV/visible wavelength emissions observed in the plumes. This emission was observed to extend several thousand meters from the shuttle, and detailed analysis shows that the total intensity of the emission depends on the ram angle (angle in the shuttle reference frame between the plume effluents and the ramming atmosphere) and altitude, indicating an interaction process with the atmosphere. Data from two observational experiments are presented. The Air Force Maui Optical Site (AMOS) experiment includes ground-based spectral and spatial measurements of the shuttle plumes as the thrusters were fired over the AMOS site on top of Haliakala Volcano on the island of Maui in the mid-Pacific. The GLO experiment was flown in the payload bay of the space shuttle and also includes spectral and spatial measurements of the shuttle plumes. During both of these experiments, the primary reaction control system (PRCS) engines (870 lb (394 kgf) thrust) and Vernier reaction control system (VRCS) engines (25 lb (11 kgf) thrust) were fired at various angles relative to the ram, thus providing a range of collision velocities (4.5-11 km/s) between the thruster plume and the atmosphere. In this report the dependence of the NH emission on ram angle, thruster size, and distance from the shuttle is presented and analyzed using a three-dimensional Monte Carlo simulation of the plume-atmosphere interactions called

  17. Development of Electrothermal Pulsed Plasma Thrusters for Osaka-Institute-of-Technology Electric-Rocket-Engine onboard Small Space Ship

    SciTech Connect

    Ishii, Yushuke; Yamamoto, Tsuyoshi; Yamada, Minetsugu; Tahara, Hirokazu

    2008-12-31

    The Project of Osaka-Institute-of-Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at Osaka Institute of Technology. In PROITERES, a 10-kg small satellite with electrothermal pulsed plasma thrusters (PPTs), named JOSHO, will be launched in 2010. The main mission is powered flight of small satellite by electric thruster itself. Electrothermal PPTs were studied with both experiments and numerical simulations. An electrothermal PPT with a side-fed propellant feeding mechanism achieved a total impulse of 3.6 Ns with a repetitive 10000-shot operation. An unsteady numerical simulation showed the existence of considerable amount of ablation delaying to the discharge. However, it was also shown that this phenomenon should not be regarded as the 'late time ablation' for electrothermal PPTs.

  18. Space Experiments with Particle Accelerators (SEPAC)

    NASA Technical Reports Server (NTRS)

    Roberts, W. T.

    1985-01-01

    The space experiments with particle accelerators (SEPAC) instruments consist of an electron accelerator, a plasma accelerator, a neutral gas (N2) release device, particle and field diagnostic instruments, and a low light level television system. These instruments are used to accomplish multiple experiments: to study beam particle interactions and other plasma processes; as probes to investigate magnetospheric processes; and as perturbation devices to study energy coupling mechanisms in the magnetosphere, ionosphere, and upper atmosphere.

  19. Space Acceleration Measurement System-II

    NASA Technical Reports Server (NTRS)

    Foster, William

    2009-01-01

    Space Acceleration Measurement System (SAMS-II) is an ongoing study of the small forces (vibrations and accelerations) on the ISS that result from the operation of hardware, crew activities, as well as dockings and maneuvering. Results will be used to generalize the types of vibrations affecting vibration-sensitive experiments. Investigators seek to better understand the vibration environment on the space station to enable future research.

  20. Delivery of Colloid Micro-Newton Thrusters for the Space Technology 7 Mission

    NASA Technical Reports Server (NTRS)

    Ziemer, John K.; Randolph, Thomas M.; Franklin, Garth W.; Hruby, Vlad; Spence, Douglas; Demmons, Nathaniel; Roy, Thomas; Ehrbar, Eric; Zwahlen, Jurg; Martin, Roy; Connolly, William

    2008-01-01

    Two flight-qualified clusters of four Colloid Micro-Newton Thruster (CMNT) systems have been delivered to the Jet Propulsion Laboratory (JPL). The clusters will provide precise spacecraft control for the drag-free technology demonstration mission, Space Technology 7 (ST7). The ST7 mission is sponsored by the NASA New Millennium Program and will demonstrate precision formation flying technologies for future missions such as the Laser Interferometer Space Antenna (LISA) mission. The ST7 disturbance reduction system (DRS) will be on the ESA LISA Pathfinder spacecraft using the European gravitational reference sensor (GRS) as part of the ESA LISA Technology Package (LTP). Developed by Busek Co. Inc., with support from JPL in design and testing, the CMNT has been developed over the last six years into a flight-ready and flight-qualified microthruster system, the first of its kind. Recent flight-unit qualification tests have included vibration and thermal vacuum environmental testing, as well as performance verification and acceptance tests. All tests have been completed successfully prior to delivery to JPL. Delivery of the first flight unit occurred in February of 2008 with the second unit following in May of 2008. Since arrival at JPL, the units have successfully passed through mass distribution, magnetic, and EMI/EMC measurements and tests as part of the integration and test (I&T) activities including the integrated avionics unit (IAU). Flight software sequences have been tested and validated with the full flight DRS instrument successfully to the extent possible in ground testing, including full functional and 72 hour autonomous operations tests. Delivery of the cluster assemblies along with the IAU to ESA for integration into the LISA Pathfinder spacecraft is planned for the summer of 2008 with a planned launch and flight demonstration in late 2010.

  1. 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.

  2. 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.

  3. Development of an iron nitrate resistant injector valve for the Space Shuttle orbiter primary thruster

    NASA Technical Reports Server (NTRS)

    Wichmann, Horst; Marquardt, Kaiser; Goforth, Alyssa

    1993-01-01

    Design of a direct-acting valve (DAV) for a primary thruster which is fully interchangeable with a thruster equipped with pilot-operated valves is described. The DAV is based on a bellows to isolate propellants form the actuator for maximum resistance to iron nitrate and other contamination and to select optimum materials for the actuator. It provides improved seal performance under all operating conditions and insensitivity to pressure transients. As compared with the existing pilot-operated valve, the DAV design is much simpler, consists of fewer parts, and will be lower in cost.

  4. Space experiments with particle accelerators

    NASA Technical Reports Server (NTRS)

    Obayashi, T.; Kawashima, N.; Kuriki, K.; Nagatomo, M.; Ninomiya, K.; Sasaki, S.; Roberts, W. T.; Chappell, C. R.; Reasoner, D. L.; Garriott, O. K.; Taylor, W. W. L.

    1984-01-01

    Electron and plasma beams and neutral gas plumes were injected into the space environment by instruuments on Spacelab 1, and various diagnostic measurements including television camera observations were performed. The results yield information on vehicle charging and neutralization, beam-plasma interactions, and ionization enhancement by neutral beam injection.

  5. 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

  6. High-Power Electromagnetic Thruster Being Developed

    NASA Technical Reports Server (NTRS)

    LaPointe, Michael R.; Mikellides, Pavlos G.

    2001-01-01

    High-power electromagnetic thrusters have been proposed as primary in-space propulsion options for several bold new interplanetary and deep-space missions. As the lead center for electric propulsion, the NASA Glenn Research Center designs, develops, and tests high-power electromagnetic technologies to meet these demanding mission requirements. Two high-power thruster concepts currently under investigation by Glenn are the magnetoplasmadynamic (MPD) thruster and the Pulsed Inductive Thruster (PIT).

  7. Particle accelerator employing transient space charge potentials

    DOEpatents

    Post, Richard F.

    1990-01-01

    The invention provides an accelerator for ions and charged particles. The plasma is generated and confined in a magnetic mirror field. The electrons of the plasma are heated to high temperatures. A series of local coils are placed along the axis of the magnetic mirror field. As an ion or particle beam is directed along the axis in sequence the coils are rapidly pulsed creating a space charge to accelerate and focus the beam of ions or charged particles.

  8. 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.

  9. 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.

  10. Causes and Mitigation of Fuel Pilot Operated Valve Pilot Seal Extrusion in Space Shuttle Orbiter Primary RCS Thrusters

    NASA Technical Reports Server (NTRS)

    Waller, Jess M.; Roth, Tim E.; Saulsberry, Regor L.; Haney, William A.; Kelly, Terence S; Forsyth, Bradley S.

    2004-01-01

    Extrusion of a polytetrafluoroethylene (PTFE) pilot seal located in the Space Shuttle Orbiter Primary Reaction Control Subsystem (PRCS) thruster fuel valve has been implicated in 68 ground and on-orbit fuel valve failures. A rash of six extrusion-related in-flight anomalies over a six-mission span from December 2001 to October 2002 led to heightened activity at various NASA centers, and the formation of a multidisciplinary team to solve the problem. Empirical and theoretical approaches were used. For example, thermomechanical analysis (TMA) and exposure tests showed that some extrusion is produced by thermal cycling; however, a review of thruster service histories did not reveal a strong link between thermal cycling and extrusion. Calculations showed that the amount of observed extrusion often exceeded the amount allowed by thermally-induced stress relief. Failure analysis of failed hardware also revealed the presence of fuel-oxidizer reaction product (FORP) inside the fuel valve pilot seal cavity, and differential scanning calorimetry (DSC) showed that the FORP was intimately associated with the pilot seal material. Component-level exposure tests showed that FORP of similar composition could be produced by adjacent oxidizer valve leakage in the absence of thruster firing. Specific gravity data showed that extruded fuel valve pilot seals were less dense than new pilot seals or oxidizer valve pilot seals, indicating permanent modification of the PTFE occurred during service. It is concluded that some thermally-induced extrusion is unavoidable; however, oxidizer leakage-induced extrusion is mostly avoidable and can be mitigated. Several engineering level mitigation strategies are discussed.

  11. Artificial Neural Network Test Support Development for the Space Shuttle PRCS Thrusters

    NASA Technical Reports Server (NTRS)

    Lehr, Mark E.

    2005-01-01

    A significant anomaly, Fuel Valve Pilot Seal Extrusion, is affecting the Shuttle Primary Reaction Control System (PRCS) Thrusters, and has caused 79 to fail. To help address this problem, a Shuttle PRCS Thruster Process Evaluation Team (TPET) was formed. The White Sands Test Facility (WSTF) and Boeing members of the TPET have identified many discrete valve current trace characteristics that are predictive of the problem. However, these are difficult and time consuming to identify and trend by manual analysis. Based on this exhaustive analysis over months, 22 thrusters previously delivered by the Depot were identified as high risk for flight failures. Although these had only recently been installed, they had to be removed from Shuttles OV103 and OV104 for reprocessing, by directive of the Shuttle Project Office. The resulting impact of the thruster removal, replacement, and valve replacement was significant (months of work and hundreds of thousands of dollars). Much of this could have been saved had the proposed Neural Network (NN) tool described in this paper been in place. In addition to the significant benefits to the Shuttle indicated above, the development and implementation of this type of testing will be the genesis for potential Quality improvements across many areas of WSTF test data analysis and will be shared with other NASA centers. Future tests can be designed to incorporate engineering experience via Artificial Neural Nets (ANN) into depot level acceptance of hardware. Additionally, results were shared with a NASA Engineering and Safety Center (NESC) Super Problem Response Team (SPRT). There was extensive interest voiced among many different personnel from several centers. There are potential spin-offs of this effort that can be directly applied to other data acquisition systems as well as vehicle health management for current and future flight vehicles.

  12. 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.

  13. Plasma thrusters from Russia

    SciTech Connect

    Lerner, E.J.

    1992-09-01

    A report on the Russian stationary plasma thrusters having plasma accelerated to high velocities by electrical and magnetic forces is described. For specific impulses of 15-20 km/sec, optimal for such applications as satellite station keeping and orbital transfer, a unit supplying 0.05 N from a 2-kW input has a 30-cm-diameter nozzle.

  14. 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.

  15. Development of Eddy Current Technique for the Detection of Stress Corrosion Cracking in Space Shuttle Primary Reaction Control Thrusters

    NASA Technical Reports Server (NTRS)

    Wincheski, Buzz; Simpson, John; Koshti, Ajay

    2006-01-01

    A recent identification of stress corrosion cracking in the Space Shuttle Primary Reaction Control System (PRCS) thrusters triggered an extensive nondestructive evaluation effort to develop techniques capable of identifying such damage on installed shuttle hardware. As a part of this effort, specially designed eddy current probes inserted into the acoustic cavity were explored for the detection of such flaws and for evaluation of the remaining material between the crack tip and acoustic cavity. The technique utilizes two orthogonal eddy current probes which are scanned under stepper motor control in the acoustic cavity to identify cracks hidden with as much as 0.060 remaining wall thickness to the cavity. As crack growth rates in this area have been determined to be very slow, such an inspection provides a large safety margin for continued operation of the critical shuttle hardware. Testing has been performed on thruster components with both actual and fabricated defects. This paper will review the design and performance of the developed eddy current inspection system. Detection of flaws as a function of remaining wall thickness will be presented along with the proposed system configuration for depot level or on-vehicle inspection capabilities.

  16. Development of Eddy Current Techniques for the Detection of Cracking in Space Shuttle Primary Reaction Control Thrusters

    NASA Technical Reports Server (NTRS)

    Wincheski, Buzz A.; Simpson, John W.; Koshti, Ajay

    2007-01-01

    A recent identification of cracking in the Space Shuttle Primary Reaction Control System (PRCS) thrusters triggered an extensive nondestructive evaluation effort to develop techniques capable of identifying such damage on installed shuttle hardware. As a part of this effort, specially designed eddy current probes inserted into the acoustic cavity were explored for the detection of such flaws and for evaluation of the remaining material between the crack tip and acoustic cavity. The technique utilizes two orthogonal eddy current probes which are scanned under stepper motor control in the acoustic cavity to identify cracks hidden with as much as 0.060 remaining wall thickness to the cavity. As crack growth rates in this area have been determined to be very slow, such an inspection provides a large safety margin for continued operation of the critical shuttle hardware. Testing has been performed on thruster components with both actual and fabricated defects. This paper will review the design and performance of the developed eddy current inspection system. Detection of flaws as a function of remaining wall thickness will be presented along with the proposed system configuration for depot level or on-vehicle inspection capabilities.

  17. 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.

  18. 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.

  19. NASA GRC High Power Electromagnetic Thruster Program

    NASA Technical Reports Server (NTRS)

    LaPointe, Michael R.; Pensil, Eric J.

    2004-01-01

    High-power electromagnetic thrusters have been proposed as primary in-space propulsion options for several bold new interplanetary and deep-space missions. As the lead center for electric propulsion, the NASA Glenn Research Center designs, develops, and tests high-power electromagnetic technologies to meet these demanding mission requirements. Two high-power thruster concepts currently under investigation by Glenn are the magnetoplasmadynamic (MPD) thruster and the Pulsed Inductive Thruster (PIT). This paper describes the MPD thruster and the test facility.

  20. 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.

  1. 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.

  2. 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.

  3. Space Experiments with Particle Accelerators: SEPAC

    NASA Technical Reports Server (NTRS)

    Burch, J. L.; Roberts, W. T.; Taylor, W. W. L.; Kawashima, N.; Marshall, J. A.; Moses, S. L.; Neubert, T.; Mende, S. B.; Choueiri, E. Y.

    1994-01-01

    The Space Experiments with Particle Accelerators (SEPAC), which flew on the Atmospheric Laboratory for Applications and Science (ATLAS) 1 mission, used new techniques to study natural phenomena in the Earth's upper atmosphere, ionosphere and magnetosphere by introducing energetic perturbations into the system from a high power electron beam with known characteristics. Properties of auroras were studied by directing the electron beam into the upper atmosphere while making measurements of optical emissions. Studies were also performed of the critical ionization velocity phenomenon.

  4. Laser-induced breakdown spectroscopy in a running Hall Effect Thruster for space propulsion

    NASA Astrophysics Data System (ADS)

    Balika, L.; Focsa, C.; Gurlui, S.; Pellerin, S.; Pellerin, N.; Pagnon, D.; Dudeck, M.

    2012-08-01

    Hall Effect Thrusters (HETs) are promising electric propulsion devices for the station-keeping of geostationary satellites and for interplanetary missions. The main limiting factor of the HET lifetime is the erosion of the annular channel ceramic walls. Erosion monitoring has been performed in the laboratory using optical emission spectroscopy (OES) measurements and data treatment based on the coronal model and the actinometric hypothesis. This study uses laser ablation of the ceramic wall in a running HET in order to introduce controlled amounts of sputtered material in the thruster plasma. The transient laser-induced breakdown plasma expands orthogonally in a steady-state plasma jet created by the HET discharge. The proposed spectroscopic method involves species from both plasmas (B, Xe, Xe+). The optical emission signal is correlated to the ablated volume (measured by profilometry) leading to the first direct validation of the actinometric hypothesis in this frame and opening the road for calibration of in-flight erosion monitoring based on the OES method.

  5. 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).

  6. 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.

  7. Characteristics of the XHT-100 Low Power Hall Thruster Prototype

    NASA Astrophysics Data System (ADS)

    Andrenucci, M.; Berti, M.; Biagioni, L.; Cesari, U.; Saverdi, M.

    2004-10-01

    Several space applications indicate the possibility to adopt Mini Hall Thrusters, with discharge power in the range 50 to 200 W, among existing electric thruster propulsion technologies, to match mission propulsion requirements. A nominally 100W Hall Effect Thruster prototype (with an alumina acceleration chamber diameter slightly larger than 29 mm) has been recently designed and manufactured by Alta and Centrospazio, with the purpose of performing a wide range parametric exploration of the main engineering and physical aspects relevant to these devices at low power. During 2004 a preliminary experimental characterization has been performed in Alta's IV-4 test facility (in Pisa, Italy), a 2 m dia. 4 m length AISI 316 L vacuum chamber, equipped with a set of 6 tailored cryopumping surfaces with a total pumping speed on Xe in the order of 70000 l/s. Additional tests will be performed at ESA- ESTEC Electric Propulsion Laboratory (in the Netherlands).

  8. 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.

  9. Space Experiments with Particle Accelerators (SEPAC)

    NASA Technical Reports Server (NTRS)

    Obayashi, Tatsuzo

    1988-01-01

    The purpose of Space Experiments with Particle Accelerators (SEPAC) on the Atmospheric Laboratory for Applications and Science (ATLAS 1) mission, is to carry out active and interactive experiments on and in the earth's ionosphere, atmosphere, and magnetosphere. The instruments to be used are an electron beam accelerator (EBA), plasma contactor, and associated instruments the purpose of which is to perform diagnostic, monitoring, and general data taking functions. Four major classes of investigations are to be performed by SEPAC. They are: beam plasma physics, beam-atmosphere interactions, the use of modulated electron beams as transmitting antennas, and the use of electron beams for remote sensing of electric and magnetic fields. The first class consists mainly of onboard plasma physics experiments to measure the effects of phenomena in the vicinity of the shuttle. The last three are concerned with remote effects and are supported by other ATLAS 1 investigations as well as by ground-based observations.

  10. 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.

  11. Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    2002-01-01

    Dr. Tom Markusic, a propulsion research engineer at the Marshall Space Flight Center (MSFC), adjusts a diagnostic laser while a pulsed plasma thruster (PPT) fires in a vacuum chamber in the background. NASA/MSFC's Propulsion Research Center (PRC) is presently investigating plasma propulsion for potential use on future nuclear-powered spacecraft missions, such as human exploration of Mars.

  12. The development of reactive fuel grains for pyrophoric relight of in-space hybrid rocket thrusters

    NASA Astrophysics Data System (ADS)

    Steiner, Matthew Wellington

    This study presents and investigates a novel hybrid fuel grain that reacts pyrophorically with gaseous oxidizer to achieve restart of a hybrid rocket motor propulsion system while reducing cost and handling concerns. This reactive fuel grain (RFG) relies on the pyrophoric nature of finely divided metal particles dispersed in a solid dicyclopentadiene (DCPD) binder, which has been shown to encapsulate air-sensitive additives until they are exposed to combustion gases. An RFG is thus effectively inert in open air in the absence of an ignition source, though the particles encapsulated within remain pyrophoric. In practice, this means that an RFG that is ignited in the vacuum of space and then extinguished will expose unoxidized pyrophoric particles, which can be used to generate sufficient heat to relight the propellant when oxidizer is flowed. The experiments outlined in this work aim to develop a suitable pyrophoric material for use in an RFG, demonstrate pyrophoric relight, and characterize performance under conditions relevant to a hybrid rocket thruster. Magnesium, lithium, calcium, and an alloy of titanium, chromium, and manganese (TiCrMn) were investigated to determine suitability of pure metals as RFG additives. Additionally, aluminum hydride (AlH3), lithium aluminum hydride (LiAlH4), lithium borohydride (LiBH4), and magnesium hydride (MgH2) were investigated to determine suitability of metals hydrides as RFG additives or as precursors for pure-metal RFG additives. Pyrophoric metals have been previously investigated as additives for increasing the regression rate of hybrid fuels, but to the author's knowledge, these materials have not been specifically investigated for their ability to ignite a propellant pyrophorically. Commercial research-grade metals were obtained as coarse powders, then ball-milled to attempt to reduce particle size below a critical diameter needed for pyrophoricity. Magnesium hydride was ball-milled and then cycled in a hydride cycling

  13. 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.

  14. Space experiments with particle accelerators. [Spacelab

    NASA Technical Reports Server (NTRS)

    Obayashi, T.

    1981-01-01

    The purpose of space experiments with particle accelerators (SEPAC) is to carry out active and interactive experiments on and in the Earth's ionosphere and magnetosphere. It is also intended to make an initial performance test for an overall program of Spacelab/SEPAC experiments. The instruments to be used are an electron beam accelerator, magnetoplasma dynamic arcjet, and associated diagnostic equipment. The accelerators are installed on the pallet, with monitoring and diagnostic observations being made by the gas plume release, beam-monitor TV, and particle-wave measuring instruments also mounted on the pallet. Command and display systems are installed in the module. Three major classes of investigations to be performed are vehicle charge neutralization, beam plasma physics, and beam atmosphere interactions. The first two are mainly onboard plasma physics experiments to measure the effect of phenomena in the vicinity of Spacelab. The last one is concerned with atmospheric modification and is supported by other Spacelab 1 investigations as well as by ground-based, remote sensing observations.

  15. 20-mN Variable Specific Impulse (Isp) Colloid Thruster

    NASA Technical Reports Server (NTRS)

    Demmons, Nathaniel

    2015-01-01

    Busek Company, Inc., has designed and manufactured an electrospray emitter capable of generating 20 mN in a compact package (7x7x1.7 in). The thruster consists of nine porous-surface emitters operating in parallel from a common propellant supply. Each emitter is capable of supporting over 70,000 electrospray emission sites with the plume from each emitter being accelerated through a single aperture, eliminating the need for individual emission site alignment to an extraction grid. The total number of emission sites during operation is expected to approach 700,000. This Phase II project optimized and characterized the thruster fabricated during the Phase I effort. Additional porous emitters also were fabricated for full-scale testing. Propellant is supplied to the thruster via existing feed-system and microvalve technology previously developed by Busek, under the NASA Space Technology 7's Disturbance Reduction System (ST7-DRS) mission and via follow-on electric propulsion programs. This project investigated methods for extending thruster life beyond the previously demonstrated 450 hours. The life-extending capabilities will be demonstrated on a subscale version of the thruster.

  16. 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.

  17. 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.

  18. Space Experiments with Particle Accelerators (SEPAC)

    NASA Technical Reports Server (NTRS)

    Obayashi, T.; Kawashima, N.; Kuriki, K.; Nagatomo, M.; Ninomiya, K.; Sasaki, S.; Ushirokawa, A.; Kudo, I.; Ejiri, M.; Roberts, W. T.

    1982-01-01

    Plans for SEPAC, an instrument array to be used on Spacelab 1 to study vehicle charging and neutralization, beam-plasma interaction in space, beam-atmospheric interaction exciting artificial aurora and airglow, and the electromagnetic-field configuration of the magnetosphere, are presented. The hardware, consisting of electron beam accelerator, magnetoplasma arcjet, neutral-gas plume generator, power supply, diagnostic package (photometer, plasma probes, particle analyzers, and plasma-wave package), TV monitor, and control and data-management unit, is described. The individual SEPAC experiments, the typical operational sequence, and the general outline of the SEPAC follow-on mission are discussed. Some of the experiments are to be joint ventures with AEPI (INS 003) and will be monitored by low-light-level TV.

  19. Space Experiments with Particle Accelerators (SEPAC)

    NASA Technical Reports Server (NTRS)

    Taylor, William W. L.

    1994-01-01

    The scientific emphasis of this contract has been on the physics of beam ionosphere interactions, in particular, what are the plasma wave levels stimulated by the Space Experiments with Particle Accelerators (SEPAC) electron beam as it is ejected from the Electron Beam Accelerator (EBA) and passes into and through the ionosphere. There were two different phenomena expected. The first was generation of plasma waves by the interaction of the DC component of the beam with the plasma of the ionosphere, by wave particle interactions. The second was the generation of waves at the pulsing frequency of the beam (AC component). This is referred to as using the beam as a virtual antenna, because the beam of electrons is a coherent electrical current confined to move along the earth's magnetic field. As in a physical antenna, a conductor at a radio or TV station, the beam virtual antenna radiates electromagnetic waves at the frequency of the current variations. These two phenomena were investigated during the period of this contract.

  20. The Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

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

    2003-01-01

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

  1. 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%.

  2. What We Have Learned By Studying The P5 Hall Thruster

    NASA Astrophysics Data System (ADS)

    Gallimore, Alec D.

    2003-05-01

    The Hall thruster is an advanced spacecraft propulsion system that uses electrical power provided by the spacecraft to generate thrust by ionizing and accelerating ions to high velocities. While Hall thrusters have been tested in laboratories for nearly forty years and first flew in space some thirty years ago, little is known about the plasma within these devices. This lack of knowledge has led to the expensive trial-and-error approach practiced in Hall thruster development over the years. The difficulty in collecting interior plasma data stems from the intense heat flux a probe receives. While optical measurements can give some information about the plasma, probes provide data about the plasma that are not accessible with optical approaches. Discharge channel plasma data are vital for extending our understanding of Hall thruster physical processes, for validating thruster models, and for developing advanced, next-generation engines for high ΔV missions. The paper summarizes the results of research aimed at using probes to characterize the internal plasma structure of a laboratory Hall thruster developed specifically for this purpose. Internal plasma parameter measurements were accomplished by using the unique High-speed Axial Reciprocating Probe (HARP) system, which enabled, for the first time, the insertion and removal of probes from the Hall thruster discharge channel while minimizing perturbation to thruster operation. The system was used with an emissive probe to map plasma potential, and a double Langmuir probe to map electron temperature and ion number density. Thruster perturbation, determined by monitoring discharge current, was less than 10% for the majority of measurements.

  3. A Microwave Thruster for Spacecraft Propulsion

    SciTech Connect

    Chiravalle, Vincent P

    2012-07-23

    This presentation describes how a microwave thruster can be used for spacecraft propulsion. A microwave thruster is part of a larger class of electric propulsion devices that have higher specific impulse and lower thrust than conventional chemical rocket engines. Examples of electric propulsion devices are given in this presentation and it is shown how these devices have been used to accomplish two recent space missions. The microwave thruster is then described and it is explained how the thrust and specific impulse of the thruster can be measured. Calculations of the gas temperature and plasma properties in the microwave thruster are discussed. In addition a potential mission for the microwave thruster involving the orbit raising of a space station is explored.

  4. 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.

  5. 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.

  6. NASA 30 Cm Ion Thruster Development Status

    NASA Technical Reports Server (NTRS)

    Patterson, Michael J.; Haag, Thomas W.; Rawlin, Vincent K.; Kussmaul, Michael T.

    1995-01-01

    A 30 cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for missions of national interest and it is an element of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) program established to validate ion propulsion for space flight applications. The thruster has been developed to an engineering model level and it incorporates innovations in design, materials, and fabrication techniques compared to those employed to conventional ion thrusters. The performance of both functional and engineering model thrusters has been assessed including thrust stand measurements, over an input power range of 0.5-2.3 kW. Attributes of the engineering model thruster include an overall mass of 6.4 kg, and an efficiency of 65 percent and thrust of 93 mN at 2.3 kW input power. This paper discusses the design, performance, and lifetime expectations of the functional and engineering model thrusters under development at NASA.

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

    NASA Astrophysics Data System (ADS)

    Rafalskyi, D.; Aanesland, A.

    2014-11-01

    We propose an alternative method to accelerate ions in classical gridded ion thrusters and ion sources such that co-extracted electrons from the source may provide beam space charge neutralization. In this way there is no need for an additional electron neutralizer. The method consists of applying RF voltage to a two-grid acceleration system via a blocking capacitor. Due to the unequal effective area of the two grids in contact with the plasma, a dc self-bias is formed, rectifying the applied RF voltage. As a result, ions are continuously accelerated within the grid system while electrons are emitted in brief instants within the RF period when the RF space charge sheath collapses. This paper presents the first experimental results and a proof-of-principle. Experiments are carried out using the Neptune thruster prototype which is a gridded Inductively Coupled Plasma (ICP) source operated at 4 MHz, attached to a larger beam propagation chamber. The RF power supply is used both for the ICP discharge (plasma generation) and powering the acceleration grids via a capacitor for ion acceleration and electron extraction without any dc power supplies. The ion and electron energies, particle flux and densities are measured using retarding field energy analyzers (RFEA), Langmuir probes and a large beam target. The system operates in Argon and N2. The dc self-bias is found to be generated within the gridded extraction system in all the range of operating conditions. Broad quasi-neutral ion-electron beams are measured in the downstream chamber with energies up to 400 eV. The beams from the RF acceleration method are compared with classical dc acceleration with an additional external electron neutralizer. It is found that the two acceleration techniques provide similar performance, but the ion energy distribution function from RF acceleration is broader, while the floating potential of the beam is lower than for the dc accelerated beam.

  8. High Power MPD Thruster Performance Measurements

    NASA Technical Reports Server (NTRS)

    LaPointe, Michael R.; Strzempkowski, Eugene; Pencil, Eric

    2004-01-01

    High power magnetoplasmadynamic (MPD) thrusters are being developed as cost effective propulsion systems for cargo transport to lunar and Mars bases, crewed missions to Mars and the outer planets, and robotic deep space exploration missions. Electromagnetic MPD thrusters have demonstrated, at the laboratory level, the ability to process megawatts of electrical power while providing significantly higher thrust densities than electrostatic 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 programs envisioned by the NASA Office of Exploration Systems, Glenn Research Center is developing and testing quasi-steady MW-class MPD thrusters as a prelude to steady state high power thruster tests. This paper provides an overview of the GRC high power pulsed thruster test facility, and presents preliminary performance data for a quasi-steady baseline MPD thruster geometry.

  9. Space experiments with particle accelerators: SEPAC

    NASA Technical Reports Server (NTRS)

    Roberts, B.

    1986-01-01

    The SEPAC instruments consist of an electron accelerator, a plasma accelerator, a neutral gas (N2) release device, particle and field diagnostic instruments, and a low light level television system. These instruments are used to accomplish multiple experiments: to study beam-particle interactions and other plasma processes; as probes to investigate magnetospheric processes; and as perturbation devices to study energy coupling mechanisms in the magnetosphere, ionosphere, and upper atmosphere.

  10. 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.

  11. Proton and heavy ion acceleration facilities for space radiation research

    NASA Technical Reports Server (NTRS)

    Miller, Jack

    2003-01-01

    The particles and energies commonly used for medium energy nuclear physics and heavy charged particle radiobiology and radiotherapy at particle accelerators are in the charge and energy range of greatest interest for space radiation health. In this article we survey some of the particle accelerator facilities in the United States and around the world that are being used for space radiation health and related research, and illustrate some of their capabilities with discussions of selected accelerator experiments applicable to the human exploration of space.

  12. Duration test of an annular colloid thruster.

    NASA Technical Reports Server (NTRS)

    Perel, J.; Mahoney, J. F.; Daley, H. L.

    1972-01-01

    An annular colloid thruster was continuously operated for 1023 hours. Performance was stable with no sparking and negligible drain currents observed. An average thrust of 25.1 micropounds and an average specific impulse of 1160 seconds were obtained at an accelerating voltage of 15 k he thruster exhaust beam was continuously neutralized using electrons and electrostatic vectoring was demonstrated periodically. The only clear trend with time was an increase in specific impulse during the last third of the test period. From these results the thruster lifetime was estimated to be over an order of magnitude greater than the test duration.

  13. Stationary plasma thruster evaluation in Russia

    NASA Technical Reports Server (NTRS)

    Brophy, John R.

    1992-01-01

    A team of electric propulsion specialists from U.S. government laboratories experimentally evaluated the performance of a 1.35-kW Stationary Plasma Thruster (SPT) at the Scientific Research Institute of Thermal Processes in Moscow and at 'Fakel' Enterprise in Kaliningrad, Russia. The evaluation was performed using a combination of U.S. and Russian instrumentation and indicated that the actual performance of the thruster appears to be close to the claimed performance. The claimed performance was a specific impulse of 16,000 m/s, an overall efficiency of 50 percent, and an input power of 1.35 kW, and is superior to the performance of western electric thrusters at this specific impulse. The unique performance capabilities of the stationary plasma thruster, along with claims that more than fifty of the 660-W thrusters have been flown in space on Russian spacecraft, attracted the interest of western spacecraft propulsion specialists. A two-phase program was initiated to evaluate the stationary plasma thruster performance and technology. The first phase of this program, to experimentally evaluate the performance of the thruster with U.S. instrumentation in Russia, is described in this report. The second phase objective is to determine the suitability of the stationary plasma thruster technology for use on western spacecraft. This will be accomplished by bringing stationary plasma thrusters to the U.S. for quantification of thruster erosion rates, measurements of the performance variation as a function of long-duration operation, quantification of the exhaust beam divergence angle, and determination of the non-propellant efflux from the thruster. These issues require quantification in order to maximize the probability for user application of the SPT technology and significantly increase the propulsion capabilities of U.S. spacecraft.

  14. Improvement of Space Shuttle Main Engine Low Frequency Acceleration Measurements

    NASA Technical Reports Server (NTRS)

    Stec, Robert C.

    1999-01-01

    The noise floor of low frequency acceleration data acquired on the Space Shuttle Main Engines is higher than desirable. Difficulties of acquiring high quality acceleration data on this engine are discussed. The approach presented in this paper for reducing the acceleration noise floor focuses on a search for an accelerometer more capable of measuring low frequency accelerations. An overview is given of the current measurement system used to acquire engine vibratory data. The severity of vibration, temperature, and moisture environments are considered. Vibratory measurements from both laboratory and rocket engine tests are presented.

  15. Presentation on a Space Acceleration Measurement System (SAMS)

    NASA Technical Reports Server (NTRS)

    Chase, Theodore L.

    1990-01-01

    The primary objective of the Space Acceleration Measurement Systems (SAMS) project is to provide an acceleration measurement system capable of serving a wide variety of space experiments. The design of the system being developed under this project takes into consideration requirements for experiments located in the middeck, in the orbiter bay, and in Spacelab. In addition to measuring, conditioning, and recording accelerations, the system will be capable of performing complex calculations and interactive control. The main components consist of a remote triaxial optical storage device. In operation, the triaxial sensor head produces output signals in response to acceleration inputs. These signals are preamplified, filtered and converted into digital data which is then transferred to optical memory. The system design is modular, facilitating both software and hardware upgrading as technology advances. Two complete acceleration measurement flight systems will be build and tested under this project.

  16. Pulsed Plasma Thruster Technology

    NASA Technical Reports Server (NTRS)

    1996-01-01

    The continuing emphasis on reducing costs and downsizing spacecraft is forcing increased emphasis on reducing the subsystem mass and integration costs. For many commercial, scientific, and Department of Defense space missions, onboard propulsion is either the predominant spacecraft mass or it limits the spacecraft lifetime. Electromagnetic-pulsed-plasma thrusters (PPT's) offer the combined benefits of extremely low average electric power requirements (1 to 150 W), high specific impulse (approx. 1000 sec), and system simplicity derived from the use of an inert solid propellant. Potential applications range from orbit insertion and maintenance of small satellites to attitude control for large geostationary communications satellites.

  17. Electrodeless Experimental Thruster

    SciTech Connect

    Brainerd, Jerome J.; Reisz, Al

    2009-03-16

    An electrodeless experimental electric thruster has been built and tested at the NASA Marshall Space Flight Center (MSFC). The plasma is formed by Electron-Cyclotron Resonance (ECR) absorption of RF waves (microwaves). The RF source operates in the 1 to 2 kW range. The plasma is overdense and is confined radially by an applied axial dc magnetic field. The field is shaped by a strong magnetic mirror on the upstream end and a magnetic nozzle on the downstream end. Argon is used as the propellant. The velocity profile in the exhaust plume has been measured with Laser Induced Fluorescence (LIF). An unusual bimodal velocity profile has been measured.

  18. Space Launch System Accelerated Booster Development Cycle

    NASA Technical Reports Server (NTRS)

    Arockiam, Nicole; Whittecar, William; Edwards, Stephen

    2012-01-01

    With the retirement of the Space Shuttle, NASA is seeking to reinvigorate the national space program and recapture the public s interest in human space exploration by developing missions to the Moon, near-earth asteroids, Lagrange points, Mars, and beyond. The would-be successor to the Space Shuttle, NASA s Constellation Program, planned to take humans back to the Moon by 2020, but due to budgetary constraints was cancelled in 2010 in search of a more "affordable, sustainable, and realistic" concept2. Following a number of studies, the much anticipated Space Launch System (SLS) was unveiled in September of 2011. The SLS core architecture consists of a cryogenic first stage with five Space Shuttle Main Engines (SSMEs), and a cryogenic second stage using a new J-2X engine3. The baseline configuration employs two 5-segment solid rocket boosters to achieve a 70 metric ton payload capability, but a new, more capable booster system will be required to attain the goal of 130 metric tons to orbit. To this end, NASA s Marshall Space Flight Center recently released a NASA Research Announcement (NRA) entitled "Space Launch System (SLS) Advanced Booster Engineering Demonstration and/or Risk Reduction." The increased emphasis on affordability is evident in the language used in the NRA, which is focused on risk reduction "leading to an affordable Advanced Booster that meets the evolved capabilities of SLS" and "enabling competition" to "enhance SLS affordability. The purpose of the work presented in this paper is to perform an independent assessment of the elements that make up an affordable and realistic path forward for the SLS booster system, utilizing advanced design methods and technology evaluation techniques. The goal is to identify elements that will enable a more sustainable development program by exploring the trade space of heavy lift booster systems and focusing on affordability, operability, and reliability at the system and subsystem levels5. For this study

  19. A particle accelerator employing transient space charge potentials

    DOEpatents

    Post, R.F.

    1988-02-25

    The invention provides an accelerator for ions and charged particles. The plasma is generated and confined in a magnetic mirror field. The electrons of the plasma are heated to high temperatures. A series of local coils are placed along the axis of the magnetic mirror field. As an ion or particle beam is directed along the axis in sequence the coils are rapidly pulsed creating a space charge to accelerate and focus the beam of ions or charged particles. 3 figs.

  20. 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.

  1. High power ion thruster performance

    NASA Technical Reports Server (NTRS)

    Rawlin, Vincent K.; Patterson, Michael J.

    1987-01-01

    The ion thruster is one of several forms of space electric propulsion being considered for use on future SP-100-based missions. One possible major mission ground rule is the use of a single Space Shuttle launch. Thus, the mass in orbit at the reactor activation altitude would be limited by the Shuttle mass constraints. When the spacecraft subsystem masses are subtracted from this available mass limit, a maximum propellant mass may be calculated. Knowing the characteristics of each type of electric thruster allows maximum values of total impulse, mission velocity increment, and thrusting time to be calculated. Because ion thrusters easily operate at high values of efficiency (60 to 70%) and specific impulse (3000 to 5000 sec), they can impart large values of total impulse to a spacecraft. They also can be operated with separate control of the propellant flow rate and exhaust velocity. This paper presents values of demonstrated and projected performance of high power ion thrusters used in an analysis of electric propulsion for an SP-100 based mission.

  2. 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.

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

    NASA Astrophysics Data System (ADS)

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

    2002-10-01

    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.

  4. 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.

  5. Effect of a Second, Parallel Capacitor on the Performance of a Pulse Inductive Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Balla, Joseph V.

    2010-01-01

    Pulsed inductive plasma accelerators are electrodeless space propulsion devices where a capacitor is charged to an initial voltage and is then discharged through an inductive coil that couples energy into the propellant, ionizing and accelerating it to produce thrust. A model that employs a set of circuit equations (as illustrated in Fig. 1a) coupled to a one-dimensional momentum equation has been previously used by Lovberg and Dailey [1] and Polzin et al. [2-4] to model the plasma acceleration process in pulsed inductive thrusters. In this paper an extra capacitor, inductor, and resistor are added to the system in the manner illustrated in the schematic shown in Fig. 1b. If the second capacitor has a smaller value than the initially charged capacitor, it can serve to increase the current rise rate through the inductive coil. Increasing the current rise rate should serve to better ionize the propellant. The equation of motion is solved to find the effect of an increased current rise rate on the acceleration process. We examine the tradeoffs between enhancing the breakdown process (increasing current rise rate) and altering the plasma acceleration process. These results provide insight into the performance of modified circuits in an inductive thruster, revealing how this design permutation can affect an inductive thruster's performance.

  6. Transient phenomena in closed electron drift plasma thrusters: insights obtained in a French cooperative program

    NASA Astrophysics Data System (ADS)

    Bouchoule, A.; Philippe-Kadlec, Ch; Prioul, M.; Darnon, F.; Lyszyk, M.; Magne, L.; Pagnon, D.; Roche, S.; Touzeau, M.; Béchu, S.; Lasgorceix, P.; Sadeghi, N.; Dorval, N.; Marque, J.-P.; Bonnet, J.

    2001-05-01

    This paper presents some aspects of the research developed in the frame of a coordinated program launched in France in 1996 and devoted to plasma thrusters for space technologies. Relevant results of physical studies have been selected from the literature with the addition of recent original results. The thrusters within the scope of this research are diagnostic equipped versions of industrial realizations, in a thrust level range of 0.1 N and electrical power 1.5 kW. The optical and electrical diagnostics concern studies of the thruster plasma and of the thruster plume. Transient phenomena in these two regions, related to discharge current fluctuations or oscillations on a typical time scale of 40 µs, have been space-time characterized. This has been achieved by developing a large panel of diagnostics including RFEA, Langmuir probes, OES, fast camera imaging and electron drift Hall current probe. They lead to a coherent representation of these phenomena , in rather good qualitative agreement with 1D modelling. But they emphasize also the importance of 2D effects. Insights obtained through combined LIF (on Xe+ ions) and OES diagnostics are also presented. They concern the ionization-acceleration region in the thruster plasma, where intrusive diagnostics are disturbing in nature, and open a new step for a significant improvement of the detailed understanding of these thrusters. Such improvements are required when looking at the final goal of a predicable modelling simulation able to help the design of optimized structures at various thrust levels, in spite of the important work devoted to these devices in the former USSR and by Russian teams in Moscow at the MIREA, MAI-RIAME and KOURCHATOV Institutes.

  7. 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.

  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. Human Outer Solar System Exploration via Q-Thruster Technology

    NASA Technical Reports Server (NTRS)

    Joosten, B. Kent; White, Harold G.

    2014-01-01

    Propulsion technology development efforts at the NASA Johnson Space Center continue to advance the understanding of the quantum vacuum plasma thruster (QThruster), a form of electric propulsion. Through the use of electric and magnetic fields, a Q-thruster pushes quantum particles (electrons/positrons) in one direction, while the Qthruster recoils to conserve momentum. This principle is similar to how a submarine uses its propeller to push water in one direction, while the submarine recoils to conserve momentum. Based on laboratory results, it appears that continuous specific thrust levels of 0.4 - 4.0 N/kWe are achievable with essentially no onboard propellant consumption. To evaluate the potential of this technology, a mission analysis tool was developed utilizing the Generalized Reduced Gradient non-linear parameter optimization engine contained in the Microsoft Excel® platform. This tool allowed very rapid assessments of "Q-Ship" minimum time transfers from earth to the outer planets and back utilizing parametric variations in thrust acceleration while enforcing constraints on planetary phase angles and minimum heliocentric distances. A conservative Q-Thruster specific thrust assumption (0.4 N/kWe) combined with "moderate" levels of space nuclear power (1 - 2 MWe) and vehicle specific mass (45 - 55 kg/kWe) results in continuous milli-g thrust acceleration, opening up realms of human spaceflight performance completely unattainable by any current systems or near-term proposed technologies. Minimum flight times to Mars are predicted to be as low as 75 days, but perhaps more importantly new "retro-phase" and "gravity-augmented" trajectory shaping techniques were revealed which overcome adverse planetary phasing and allow virtually unrestricted departure and return opportunities. Even more impressively, the Jovian and Saturnian systems would be opened up to human exploration with round-trip times of 21 and 32 months respectively including 6 to 12 months of

  10. Space experiments with particle accelerators (SEPAC): Description of instrumentation

    NASA Technical Reports Server (NTRS)

    Taylor, W. W. L.; Roberts, W. T.; Reasoner, D. L.; Chappell, C. R.; Baker, B. B.; Burch, J. L.; Gibson, W. C.; Black, R. K.; Tomlinson, W. M.; Bounds, J. R.

    1987-01-01

    SEPAC (Space Experiments with Particle Accelerators) flew on Spacelab 1 (SL 1) in November and December 1983. SEPAC is a joint U.S.-Japan investigation of the interaction of electron, plasma, and neutral beams with the ionosphere, atmosphere and magnetosphere. It is scheduled to fly again on Atlas 1 in August 1990. On SL 1, SEPAC used an electron accelerator, a plasma accelerator, and neutral gas source as active elements and an array of diagnostics to investigate the interactions. For Atlas 1, the plasma accelerator will be replaced by a plasma contactor and charge collection devices to improve vehicle charging meutralization. This paper describes the SEPAC instrumentation in detail for the SL 1 and Atlas 1 flights and includes a bibliography of SEPAC papers.

  11. A multiple-cathode, high-power, rectangular ion thruster discharge chamber of increasing thruster lifetime

    NASA Astrophysics Data System (ADS)

    Rovey, Joshua Lucas

    Ion thrusters are high-efficiency, high-specific impulse space propulsion systems proposed for deep space missions requiring thruster operational lifetimes of 7--14 years. One of the primary ion thruster components is the discharge cathode assembly (DCA). The DCA initiates and sustains ion thruster operation. Contemporary ion thrusters utilize one molybdenum keeper DCA that lasts only ˜30,000 hours (˜3 years), so single-DCA ion thrusters are incapable of satisfying the mission requirements. The aim of this work is to develop an ion thruster that sequentially operates multiple DCAs to increase thruster lifetime. If a single-DCA ion thruster can operate 3 years, then perhaps a triple-DCA thruster can operate 9 years. Initially, a multiple-cathode discharge chamber (MCDC) is designed and fabricated. Performance curves and grid-plane current uniformity indicate operation similar to other thrusters. Specifically, the configuration that balances both performance and uniformity provides a production cost of 194 W/A at 89% propellant efficiency with a flatness parameter of 0.55. One of the primary MCDC concerns is the effect an operating DCA has on the two dormant cathodes. Multiple experiments are conducted to determine plasma properties throughout the MCDC and near the dormant cathodes, including using "dummy" cathodes outfitted with plasma diagnostics and internal plasma property mapping. Results are utilized in an erosion analysis that suggests dormant cathodes suffer a maximum pre-operation erosion rate of 5--15 mum/khr (active DCA maximum erosion is 70 mum/khr). Lifetime predictions indicate that triple-DCA MCDC lifetime is approximately 2.5 times longer than a single-DCA thruster. Also, utilization of new keeper materials, such as carbon graphite, may significantly decrease both active and dormant cathode erosion, leading to a further increase in thruster lifetime. Finally, a theory based on the near-DCA plasma potential structure and propellant flow rate effects

  12. 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.

  13. 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.

  14. Laser-heated rocket thruster

    NASA Technical Reports Server (NTRS)

    Shoji, J. M.

    1977-01-01

    A space vehicle application using 5,000-kw input laser power was conceptually evaluated. A detailed design evaluation of a 10-kw experimental thruster including plasma size, chamber size, cooling, and performance analyses, was performed for 50 psia chamber pressure and using hydrogen as a propellant. The 10-kw hardware fabricated included a water cooled chamber, an uncooled copper chamber, an injector, igniters, and a thrust stand. A 10-kw optical train was designed.

  15. 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.

  16. 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.

  17. Direct longitudinal laser acceleration of electrons in free space

    NASA Astrophysics Data System (ADS)

    Carbajo, Sergio; Nanni, Emilio A.; Wong, Liang Jie; Moriena, Gustavo; Keathley, Phillip D.; Laurent, Guillaume; Miller, R. J. Dwayne; Kärtner, Franz X.

    2016-02-01

    Compact laser-driven accelerators are pursued heavily worldwide because they make novel methods and tools invented at national laboratories widely accessible in science, health, security, and technology [V. Malka et al., Principles and applications of compact laser-plasma accelerators, Nat. Phys. 4, 447 (2008)]. Current leading laser-based accelerator technologies [S. P. D. Mangles et al., Monoenergetic beams of relativistic electrons from intense laser-plasma interactions, Nature (London) 431, 535 (2004); T. Toncian et al., Ultrafast laser-driven microlens to focus and energy-select mega-electron volt protons, Science 312, 410 (2006); S. Tokita et al. Single-shot ultrafast electron diffraction with a laser-accelerated sub-MeV electron pulse, Appl. Phys. Lett. 95, 111911 (2009)] rely on a medium to assist the light to particle energy transfer. The medium imposes material limitations or may introduce inhomogeneous fields [J. R. Dwyer et al., Femtosecond electron diffraction: "Making the molecular movie,", Phil. Trans. R. Soc. A 364, 741 (2006)]. The advent of few cycle ultraintense radially polarized lasers [S. Carbajo et al., Efficient generation of ultraintense few-cycle radially polarized laser pulses, Opt. Lett. 39, 2487 (2014)] has ushered in a novel accelerator concept [L. J. Wong and F. X. Kärtner, Direct acceleration of an electron in infinite vacuum by a pulsed radially polarized laser beam, Opt. Express 18, 25035 (2010); F. Pierre-Louis et al. Direct-field electron acceleration with ultrafast radially polarized laser beams: Scaling laws and optimization, J. Phys. B 43, 025401 (2010); Y. I. Salamin, Electron acceleration from rest in vacuum by an axicon Gaussian laser beam, Phys. Rev. A 73, 043402 (2006); C. Varin and M. Piché, Relativistic attosecond electron pulses from a free-space laser-acceleration scheme, Phys. Rev. E 74, 045602 (2006); A. Sell and F. X. Kärtner, Attosecond electron bunches accelerated and compressed by radially polarized laser

  18. 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.

  19. 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.

  20. Helicon Plasma Injection into an IEC Thruster

    NASA Astrophysics Data System (ADS)

    Miley, George H.; Ulmen, Ben; Krishnamurthy, Akshata; Keutelian, Paul; Chen, George

    2012-10-01

    Helicon plasma injection into an Inertial Electrostatic Confinement (IEC) thruster stage is under experimental study. Helicons are RF plasma sources using helicon waves, or low frequency whistler waves. Such inductively coupled plasmas (ICPs) produce plasma of higher density than other field-free ICPs. Permanent magnet helicon sources have been proposed for plasma generation to provide a high downstream plasma density of about 1018 m-3 for argon [1]. The IEC stage then provides plasma acceleration. An IEC plasma is produced as a dc glow discharge, giving a ``star'' or a ``jet'' mode (current mode of interest) on chamber pressure. To decouple the chamber pressure and the IEC cathode grid accelerating voltage, the plasma generation and plasma acceleration stages are decoupled by helicon plasma injection. Normally a symmetric cathode grid produces the star mode at lower pressures while a slightly higher pressure gives the jet mode. In this experiment an asymmetric cathode grid is used in the IEC to produce a plasma jet at pressures as low as 1.0 mTorr. This mode can be used as an advanced electric propulsion system where thrust and specific impulse are decoupled, providing variable specific impulse that enables complicated orbital maneuvers and challenging space missions. [4pt] [1] F. F. Chen and H. Torreblanca, ``Large-area helicon plasma source with permanent magnets,'' Phys. Plasmas, 2007.

  1. 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.

  2. Ram accelerator direct space launch system - New concepts

    NASA Technical Reports Server (NTRS)

    Bogdanoff, David W.

    1992-01-01

    The ram accelerator, a chemically driven ramjet-in-tube device is a new option for direct launch of acceleration-insensitive payloads into earth orbit. The projectile is the centerbody of a ramjet and travels through a tube filled with a premixed fuel-oxidizer mixture. The tube acts as the cowl of the ramjet. A number of new concepts for a ram accelerator space launch system are presented. The velocity and acceleration capabilities of a number of ram accelerator drive modes, including several new modes, are given. Passive (fin) stabilization during atmospheric transit is investigated and found to be promising. Gasdynamic heating in-tube and during atmospheric transit is studied; the former is found to be severe, but may be alleviated by the selection of the most suitable drive modes, transpiration cooling, or a hydrogen gas core in the launch tube. To place the payload in earth orbit, scenarios using one impulse and three impulses (with an aeropass) and a new scenario involving an auxiliary vehicle are studied. The auxiliary vehicle scenario is found to be competitive regarding payload, and requires a much simpler projectile, but has the disadvantage of requiring the auxiliary vehicle.

  3. Investigation of the Effects of Facility Background Pressure on the Performance and Voltage-Current Characteristics of the High Voltage Hall Accelerator

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Spektor, Rostislav

    2014-01-01

    The National Aeronautics and Space Administration (NASA) Science Mission Directorate In-Space Propulsion Technology office is sponsoring NASA Glenn Research Center to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. A study was conducted to assess the impact of varying the facility background pressure on the High Voltage Hall Accelerator (HiVHAc) thruster performance and voltage-current characteristics. This present study evaluated the HiVHAc thruster performance in the lowest attainable background pressure condition at NASA GRC Vacuum Facility 5 to best simulate space-like conditions. Additional tests were performed at selected thruster operating conditions to investigate and elucidate the underlying physics that change during thruster operation at elevated facility background pressure. Tests were performed at background pressure conditions that are three and ten times higher than the lowest realized background pressure. Results indicated that the thruster discharge specific impulse and efficiency increased with elevated facility background pressure. The voltage-current profiles indicated a narrower stable operating region with increased background pressure. Experimental observations of the thruster operation indicated that increasing the facility background pressure shifted the ionization and acceleration zones upstream towards the thruster's anode. Future tests of the HiVHAc thruster are planned at background pressure conditions that are expected to be two to three times lower than what was achieved during this test campaign. These tests will not only assess the impact of reduced facility background pressure on thruster performance, voltage-current characteristics, and plume properties; but will also attempt to quantify the magnitude of the ionization and acceleration zones upstream shifting as a function of increased background pressure.

  4. Late time solution for interacting scalar in accelerating spaces

    SciTech Connect

    Prokopec, Tomislav

    2015-11-01

    We consider stochastic inflation in an interacting scalar field in spatially homogeneous accelerating space-times with a constant principal slow roll parameter ε. We show that, if the scalar potential is scale invariant (which is the case when scalar contains quartic self-interaction and couples non-minimally to gravity), the late-time solution on accelerating FLRW spaces can be described by a probability distribution function (PDF) ρ which is a function of φ/H only, where φ=φ( x-vector ) is the scalar field and H=H(t) denotes the Hubble parameter. We give explicit late-time solutions for ρarrow ρ{sub ∞}(φ/H), and thereby find the order ε corrections to the Starobinsky-Yokoyama result. This PDF can then be used to calculate e.g. various n-point functions of the (self-interacting) scalar field, which are valid at late times in arbitrary accelerating space-times with ε= constant.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. Requirements for Simulating Space Radiation With Particle Accelerators

    NASA Technical Reports Server (NTRS)

    Schimmerling, W.; Wilson, J. W.; Cucinotta, F.; Kim, M-H Y.

    2004-01-01

    Interplanetary space radiation consists of fully ionized nuclei of atomic elements with high energy for which only the few lowest energy ions can be stopped in shielding materials. The health risk from exposure to these ions and their secondary radiations generated in the materials of spacecraft and planetary surface enclosures is a major limiting factor in the management of space radiation risk. Accurate risk prediction depends on a knowledge of basic radiobiological mechanisms and how they are modified in the living tissues of a whole organism. To a large extent, this knowledge is not currently available. It is best developed at ground-based laboratories, using particle accelerator beams to simulate the components of space radiation. Different particles, in different energy regions, are required to study different biological effects, including beams of argon and iron nuclei in the energy range 600 to several thousand MeV/nucleon and carbon beams in the energy range of approximately 100 MeV/nucleon to approximately 1000 MeV/nucleon. Three facilities, one each in the United States, in Germany and in Japan, currently have the partial capability to satisfy these constraints. A facility has been proposed using the Brookhaven National Laboratory Booster Synchrotron in the United States; in conjunction with other on-site accelerators, it will be able to provide the full range of heavy ion beams and energies required. International cooperation in the use of these facilities is essential to the development of a safe international space program.

  10. Comparison of Computed and Measured Performance of a Pulsed Inductive Thruster Operating on Argon Propellant

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Sankaran, Kameshwaran; Ritchie, Andrew G.; Peneau, Jarred P.

    2012-01-01

    Pulsed inductive plasma accelerators are electrodeless space propulsion devices where a capacitor is charged to an initial voltage and then discharged through a coil as a high-current pulse that inductively couples energy into the propellant. The field produced by this pulse ionizes the propellant, producing a plasma near the face of the coil. Once a plasma is formed if can be accelerated and expelled at a high exhaust velocity by the Lorentz force arising from the interaction of an induced plasma current and the magnetic field. A recent review of the developmental history of planar-geometry pulsed inductive thrusters, where the coil take the shape of a flat spiral, can be found in Ref. [1]. Two concepts that have employed this geometry are the Pulsed Inductive Thruster (PIT)[2, 3] and the Faraday Accelerator with Radio-frequency Assisted Discharge (FARAD)[4]. There exists a 1-D pulsed inductive acceleration model that employs a set of circuit equations coupled to a one-dimensional momentum equation. The model was originally developed and used by Lovberg and Dailey[2, 3] and has since been nondimensionalized and used by Polzin et al.[5, 6] to define a set of scaling parameters and gain general insight into their effect on thruster performance. The circuit presented in Fig. 1 provides a description of the electrical coupling between the current flowing in the thruster I1 and the plasma current I2. Recently, the model was upgraded to include an equation governing the deposition of energy into various modes present in a pulsed inductive thruster system (acceleration, magnetic flux generation, resistive heating, etc.)[7]. An MHD description of the plasma energy density evolution was tailored to the thruster geometry by assuming only one-dimensional motion and averaging the plasma properties over the spatial dimensions of the current sheet to obtain an equation for the time-evolution of the total energy. The equation set governing the dynamics of the coupled

  11. Performance and Environmental Test Results of the High Voltage Hall Accelerator Engineering Development Unit

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hani; Haag, Thomas; Huang, Wensheng; Shastry, Rohit; Pinero, Luis; Peterson, Todd; Mathers, Alex

    2012-01-01

    NASA Science Mission Directorate's In-Space Propulsion Technology Program is sponsoring the development of a 3.5 kW-class engineering development unit Hall thruster for implementation in NASA science and exploration missions. NASA Glenn and Aerojet are developing a high fidelity high voltage Hall accelerator that can achieve specific impulse magnitudes greater than 2,700 seconds and xenon throughput capability in excess of 300 kilograms. Performance, plume mappings, thermal characterization, and vibration tests of the high voltage Hall accelerator engineering development unit have been performed. Performance test results indicated that at 3.9 kW the thruster achieved a total thrust efficiency and specific impulse of 58%, and 2,700 sec, respectively. Thermal characterization tests indicated that the thruster component temperatures were within the prescribed material maximum operating temperature limits during full power thruster operation. Finally, thruster vibration tests indicated that the thruster survived the 3-axes qualification full-level random vibration test series. Pre and post-vibration test performance mappings indicated almost identical thruster performance. Finally, an update on the development progress of a power processing unit and a xenon feed system is provided.

  12. 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.

  13. 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.

  14. 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.

  15. 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.

  16. Space Acceleration Measurement System (SAMS)/Orbital Acceleration Research Experiment (OARE)

    NASA Technical Reports Server (NTRS)

    Hakimzadeh, Roshanak

    1998-01-01

    The Life and Microgravity Spacelab (LMS) payload flew on the Orbiter Columbia on mission STS-78 from June 20th to July 7th, 1996. The LMS payload on STS-78 was dedicated to life sciences and microgravity experiments. Two accelerometer systems managed by the NASA Lewis Research Center (LERC) flew to support these experiments, namely the Orbital Acceleration Research Experiment (OARE) and the Space Acceleration Measurements System (SAMS). In addition, the Microgravity Measurement Assembly (NOAA), managed by the European Space Research and Technology Center (ESA/ESTEC), and sponsored by NASA, collected acceleration data in support of the experiments on-board the LMS mission. OARE downlinked real-time quasi-steady acceleration data, which was provided to the investigators. The SAMS recorded higher frequency data on-board for post-mission analysis. The MMA downlinked real-time quasi-steady as well as higher frequency acceleration data, which was provided to the investigators. The Principal Investigator Microgravity Services (PIMS) project at NASA LERC supports principal investigators of microgravity experiments as they evaluate the effects of varying acceleration levels on their experiments. A summary report was prepared by PIMS to furnish interested experiment investigators with a guide to evaluate the acceleration environment during STS-78, and as a means of identifying areas which require further study. The summary report provides an overview of the STS-78 mission, describes the accelerometer systems flown on this mission, discusses some specific analyses of the accelerometer data in relation to the various activities which occurred during the mission, and presents plots resulting from these analyses as a snapshot of the environment during the mission. Numerous activities occurred during the STS-78 mission that are of interest to the low-gravity community. Specific activities of interest during this mission were crew exercise, radiator deployment, Vernier Reaction

  17. Evaluation of a steady state MPD thruster test facility

    SciTech Connect

    Reed, C.B.; Carlson, L.W.; Herman, H.; Doss, E.D.; Kilgore, O.

    1985-01-01

    The successful development of multimegawatt MPD thrusters depends, to a great extent, on testing them under steady state high altitude space conditions. Steady state testing is required to provide thermal characteristics, life cycle, erosion, and other essential data. the major technical obstacle for ground testing of MPD thrusters in a space simulation facility is the inability of state-of-the-art vacuum systems to handle the tremendous pumping speeds required for multimegawatt MPD thrusters. This is true for other types of electric propulsion devices as well. This paper discusses the results of the first phase of an evaluation of steady state MPD thruster test facilities. The first phase addresses the conceptual design of vacuum systems required to support multimegawatt MPD thruster testing. Three advanced pumping system concepts were evaluated and are presented here.

  18. Liquid-Metal-Fed Pulsed Plasma Thrusters

    NASA Technical Reports Server (NTRS)

    Markusic, Thomas

    2004-01-01

    A short document proposes liquid-metal-fed pulsed plasma thrusters for small spacecraft. The propellant liquid for such a thruster would be a low-melting-temperature metal that would be stored molten in an unpressurized, heated reservoir and would be pumped to the thruster by a magnetohydrodynamic coupler. The liquid would enter the thruster via a metal tube inside an electrically insulating ceramic tube. A capacitor would be connected between the outlet of the metal tube and the outer electrode of the thruster. The pumping would cause a drop of liquid to form at the outlet, eventually growing large enough to make contact with the outer electrode. Contact would close the circuit through the capacitor, causing the capacitor to discharge through the drop. The capacitor would have been charged with enough energy that the discharge would vaporize, ionize, and electromagnetically accelerate the contents of the metal drop. The resulting plasma would be ejected at a speed of about 50 km/s. The vaporization of the drop would reopen the circuit through the capacitor, enabling recharging of the capacitor. As pumping continued, a new drop would grow and the process would repeat.

  19. Ram accelerator direct launch system for space cargo

    NASA Technical Reports Server (NTRS)

    1987-01-01

    A new method of efficiently accelerating relatively large masses (up to several metric tons) to velocities of 0.6 km/sec up to 12 km/sec using chemical energy has been developed. The vehicle travels through a tube filled with a premixed gaseous fuel and oxidizer mixture. There is no propellant on-board the vehicle. The tube acts as the outer cowling of a ram jet and the energy release process travels with the vehicle. The ballistic efficiency remains high up to extremely high velocities and the acceleration can be maintained at a nearly constant level. Five modes of ram accelerator operation have been investigated; these modes differ primarily in the method of chemical heat release and the operational velocity range, and include two subsonic combustion modes (one of which involves thermally choke a combustion behind the vehicle) and three detonation drive modes. These modes of propulsion are capable of efficient acceleration in the range of 0.6-12 km/sec, although aerodynamic heating becomes severe above about 8 km/sec. Experiments carried out to date at the University of Washington up to 2 km/sec have established proof of principle of the ram accelerator concept and have shown close agreement between predicted and measured performance. A launch system capable of delivering two metric tons into low earth orbit was selected for the purposes of the present study. The preliminary analysis indicates that the overall dimensions of a restricted acceleration (less than approx. 1000 g) launch facility would require a tube 1 m in diameter, with an overall length of approximately 4 km. As in any direct launch scheme, a small on-board rocket is required to circularize the otherwise highly elliptical orbit which intersects the Earth. Various orbital insertion scenarios have been explored for the case of a 9 km/sec ram accelerator launch. These include direct insertion through a single circularization maneuver (i.e., on rocket burn), insertion involving two burns, and a

  20. Investigation of the Effects of Facility Background Pressure on the Performance and Voltage-Current Characteristics of the High Voltage Hall Accelerator

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Spektor, Rostislav

    2014-01-01

    The National Aeronautics and Space Administration (NASA) Science Mission Directorate In-Space Propulsion Technology office is sponsoring NASA Glenn Research Center to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. A study was conducted to assess the impact of varying the facility background pressure on the High Voltage Hall Accelerator (HiVHAc) thruster performance and voltage-current characteristics. This present study evaluated the HiVHAc thruster performance in the lowest attainable background pressure condition at NASA GRC Vacuum Facility 5 to best simulate space-like conditions. Additional tests were performed at selected thruster operating conditions to investigate and elucidate the underlying physics that change during thruster operation at elevated facility background pressure. Tests were performed at background pressure conditions that are three and ten times higher than the lowest realized background pressure. Results indicated that the thruster discharge specific impulse and efficiency increased with elevated facility background pressure. The voltage-current profiles indicated a narrower stable operating region with increased background pressure. Experimental observations of the thruster operation indicated that increasing the facility background pressure shifted the ionization and acceleration zones upstream towards the thrusters anode. Future tests of the HiVHAc thruster are planned at background pressure conditions that are expected to be two to three times lower than what was achieved during this test campaign. These tests will not only assess the impact of reduced facility background pressure on thruster performance, voltage-current characteristics, and plume properties; but will also attempt to quantify the magnitude of the ionization.

  1. Some Considerations on the Pulsed Electromagnetic Acceleration of Plasma

    NASA Technical Reports Server (NTRS)

    Thio, Y. C. F.; Markusic, T. E.; Cassibry, J. T.; Sommers, J. C.; Turchi, P. J.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    In applying pulsed electromagnetic acceleration of plasma to space propulsion (known as pulsed plasma thrusters in the community), the mode of acceleration used has been mostly in the collisionless or near-collisionless regime. The preparation of the initial plasma is given scant attention. Collisional regime of accelerating the plasma, however, have been encountered in a variety of plasma accelerating devices. Both of these modes of acceleration are reviewed in a companion paper. In this paper, we discuss the considerations governing the controlled introduction and preparation of the initial plasma, so that the collisional mode of accelerating the plasma may be suitably enhanced.

  2. Gimballing Spacecraft Thruster

    NASA Technical Reports Server (NTRS)

    Pickens, Tim; Bossard, John

    2010-01-01

    A gimballing spacecraft reaction-control-system thruster was developed that consists of a small hydrogen/oxygen-burning rocket engine integrated with a Canfield joint. (Named after its inventor, a Canfield joint is a special gimbal mount that is strong and stable yet allows a wide range of motion.) One especially notable aspect of the design of this thruster is integration, into both the stationary legs and the moving arms of the Canfield joint, of the passages through which the hydrogen and oxygen flow to the engine. The thruster was assembled and subjected to tests in which the engine was successfully fired both with and without motion in the Canfield joint.

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

    SciTech Connect

    Cannat, F. E-mail: felix.cannat@gmail.com; Lafleur, T.; Jarrige, J.; Elias, P.-Q.; Packan, D.; Chabert, P.

    2015-05-15

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

  4. 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.

  5. 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.

  6. 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.

  7. 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.

  8. 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.

  9. Pulsed Plasma Thruster (PPT) Technology: Earth Observing-1 PPT Operational and Advanced Components Being Developed

    NASA Technical Reports Server (NTRS)

    Pencil, Eric J.; Benson, Scott W.; Arrington, Lynn A.; Frus, John; Hoskins, W. Andrew; Burton, Rodney

    2003-01-01

    In 2002 the pulsed plasma thruster (PPT) mounted on the Earth Observing-1 spacecraft was operated successfully in orbit. The two-axis thruster system is fully incorporated in the attitude determination and control system and is being used to automatically counteract disturbances in the pitch axis of the spacecraft. The first tests conducted in space demonstrated the full range of PPT operation, followed by calibration of control torques from the PPT in the attitude control system. Then the spacecraft was placed in PPT control mode. To date, it has operated for about 30 hr. The PPT successfully controlled pitch momentum during wheel de-spin, solar array acceleration and deceleration during array rewind, and environmental torques in nominal operating conditions. Images collected with the Advanced Landsat Imager during PPT operation have demonstrated that there was no degradation in comparison to full momentum wheel control. In addition, other experiments have been performed to interrogate the effects of PPT operation on communication packages and light reflection from spacecraft surfaces. Future experiments will investigate the possibility of orbit-raising maneuvers, spacecraft roll, and concurrent operation with the Hyperion imager. Future applications envisioned for pulsed plasma thrusters include longer life, higher precision, multiaxis thruster configurations for three-axis attitude control systems or high-precision, formationflying systems. Advanced components, such as a "dry" mica-foil capacitor, a wear-resistant spark plug, and a multichannel power processing unit have been developed under contract with Unison Industries, General Dynamics, and C.U. Aerospace. Over the last year, evaluation tests have been conducted to determine power processing unit efficiency, atmospheric functionality, vacuum functionality, thruster performance evaluation, thermal performance, and component life.

  10. 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.

  11. Fluid physics phenomena of resistojet thrusters

    NASA Technical Reports Server (NTRS)

    DeWitt, Kenneth J. (Principal Investigator)

    1996-01-01

    This final report includes a list of publications and part of an M.S. thesis titled 'Analyses in Theoretical and Experimental Fluid Flow', by Tony G. Howell. The thesis discusses analyses of momentum and heat transfer occurring in a laminar boundary layer of a non-Newtonian power-law fluid, and experiments completed in a simulated space thruster's plume for prediction comparison.

  12. Ion thruster design and analysis

    NASA Technical Reports Server (NTRS)

    Kami, S.; Schnelker, D. E.

    1976-01-01

    Questions concerning the mechanical design of a thruster are considered, taking into account differences in the design of an 8-cm and a 30-cm model. The components of a thruster include the thruster shell assembly, the ion extraction electrode assembly, the cathode isolator vaporizer assembly, the neutralizer isolator vaporizer assembly, ground screen and mask, and the main isolator vaporizer assembly. Attention is given to the materials used in thruster fabrication, the advanced manufacturing methods used, details of thruster performance, an evaluation of thruster life, structural and thermal design considerations, and questions of reliability and quality assurance.

  13. 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.

  14. PREFACE: Acceleration and radiation generation in space and laboratory plasmas

    NASA Astrophysics Data System (ADS)

    Bingham, R.; Katsouleas, T.; Dawson, J. M.; Stenflo, L.

    1994-01-01

    Sixty-six leading researchers from ten nations gathered in the Homeric village of Kardamyli, on the southern coast of mainland Greece, from August 29-September 4, 1993 for the International Workshop on Acceleration and Radiation Generation in Space and Laboratory Plasmas. This Special Issue represents a cross-section of the presentations made at and the research stimulated by that meeting. According to the Iliad, King Agamemnon used Kardamyli as a dowry offering in order to draw a sulking Achilles into the Trojan War. 3000 years later, Kardamyli is no less seductive. Its remoteness and tranquility made it an ideal venue for promoting the free exchange of ideas between various disciplines that do not normally interact. Through invited presen tations, informal poster discussions and working group sessions, the Workshop brought together leaders from the laboratory and space/astrophysics communities working on common problems of acceleration and radiation generation in plasmas. It was clear from the presentation and discussion sessions that there is a great deal of common ground between these disciplines which is not at first obvious due to the differing terminologies and types of observations available to each community. All of the papers in this Special Issue highlight the role collective plasma processes play in accelerating particles or generating radiation. Some are state-of-the-art presentations of the latest research in a single discipline, while others investi gate the applicability of known laboratory mechanisms to explain observations in natural plasmas. Notable among the latter are the papers by Marshall et al. on kHz radiation in the magnetosphere ; Barletta et al. on collective acceleration in solar flares; and by Dendy et al. on ion cyclotron emission. The papers in this Issue are organized as follows: In Section 1 are four general papers by Dawson, Galeev, Bingham et al. and Mon which serves as an introduction to the physical mechanisms of acceleration

  15. 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.

  16. 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

  17. Magnetic field and quadruple Langmuir probe measurements in the plume of the plasmoid thruster experiment

    NASA Astrophysics Data System (ADS)

    Koelfgen, Syri Jo

    The development of high specific impulse rocket engines is essential for fast and efficient space travel. The plasmoid thruster, a novel propulsion concept with the potential for producing a high specific impulse, was investigated in light of this need. This pulsed inductive rocket utilizes the Lorentz force to accelerate plasmoids and produce thrust. The Plasmoid Thruster Experiment (PTX) was designed to experimentally evaluate the thruster concept. PTX operates by producing plasmoids in a conical theta-pinch coil and ejecting them at high velocity. Measurements of the plasmoid magnetic fields, electron temperature (Te), electron number density (n e) and Mach number (M) were taken in the PTX plume with a B˙ probe array and a quadruple Langmuir probe. The measurements were used for calculating exit velocity and Isp. High-speed photographs were also obtained for capturing images of the plasmoids and estimating their velocity. The magnetic field data showed behavior characteristic of plasmoids, such as the occurrence of the maximum axial magnetic field on axis and magnetic field reversal. The quadruple Langmuir probe data revealed several factors that influence thruster operation, including propellant choice, supply pressure and propellant injection timing (tpuff). For Ar propellant at supply pressures of 14--34 psig and tpuff = 2200 mus, Te ranged from 2--7 eV, ne ranged from 1.5 x 1020 m-3 to 3.5 x 1020 m-3, and M ranged from 3.3--3.8 in PTX. For H2 propellant, T e ranged from 15--27 eV, ne ranged from 0.8 x 1020 m-3 to 1.5 x 1020 m-3, and M ranged from 1.4--2.6, for supply pressures of 9--38 psig and tpuff = 1200--2400 mus. Analysis of the plume measurements yielded high thruster exit velocities, indicating that the plasmoid thruster can produce a high Isp. Velocities of 24 km/s, 35 km/s and 46 km/s were calculated for supply pressures of 38 psig, 24 psig and 9 psig of H2 propellant, respectively. These exit velocities deliver Isp values of 2,400 s, 3,500 s and 4

  18. The Awful Truth About Zero-Gravity: Space Acceleration Measurement System; Orbital Acceleration Research Experiment

    NASA Technical Reports Server (NTRS)

    2002-01-01

    Earth's gravity holds the Shuttle in orbit, as it does satellites and the Moon. The apparent weightlessness experienced by astronauts and experiments on the Shuttle is a balancing act, the result of free-fall, or continuously falling around Earth. An easy way to visualize what is happening is with a thought experiment that Sir Isaac Newton did in 1686. Newton envisioned a mountain extending above Earth's atmosphere so that friction with the air would be eliminated. He imagined a cannon atop the mountain and aimed parallel to the ground. Firing the cannon propels the cannonball forward. At the same time, Earth's gravity pulls the cannonball down to the surface and eventual impact. Newton visualized using enough powder to just balance gravity so the cannonball would circle the Earth. Like the cannonball, objects orbiting Earth are in continuous free-fall, and it appears that gravity has been eliminated. Yet, that appearance is deceiving. Activities aboard the Shuttle generate a range of accelerations that have effects similar to those of gravity. The crew works and exercises. The main data relay antenna quivers 17 times per second to prevent 'stiction,' where parts stick then release with a jerk. Cooling pumps, air fans, and other systems add vibration. And traces of Earth's atmosphere, even 200 miles up, drag on the Shuttle. While imperceptible to us, these vibrations can have a profound impact on the commercial research and scientific experiments aboard the Shuttle. Measuring these forces is necessary so that researchers and scientists can see what may have affected their experiments when analyzing data. On STS-107 this service is provided by the Space Acceleration Measurement System for Free Flyers (SAMS-FF) and the Orbital Acceleration Research Experiment (OARE). Precision data from these two instruments will help scientists analyze data from their experiments and eliminate outside influences from the phenomena they are studying during the mission.

  19. Galium Electromagnetic (GEM) Thruster Concept and Design

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Markusic, Thomas E.

    2005-01-01

    We describe the design of a new type of two-stage pulsed electromagnetic accelerator, the gallium electromagnetic (GEM) thruster. A schematic illustration of the GEM thruster concept is given. In this concept, liquid gallium propellant is pumped into the first stage through a porous metal electrode using an electromagnetic pump. At a designated time, a pulsed discharge (approx. 10-50 J) is initiated in the first stage, ablating the liquid gallium from the porous electrode surface and ejecting a dense thermal gallium plasma into the second state. The presence of the gallium plasma in the second stage serves to trigger the high-energy (approx. 500 J), second-stage pulse which provides the primary electromagnetic (j x B) acceleration.

  20. 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.

  1. 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.

  2. 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.

  3. Hollow cathode and thruster discharge chamber plasma measurements

    NASA Technical Reports Server (NTRS)

    Jameson, Kristina K.; Goebel, Dan M.; Watkins, Ron M.

    2005-01-01

    Due to the successful performance of the NSTAR ion thruster in Deep Space 1 mission, coupled with the recently completed 30,352 hour extended life test (ELT) of the NSTAR flight spare thruster, ion thrusters have become a viable option for future NASA missions. In this paper, detailed measurements of the plasma parameters internal and external to the cathode will presented for the NSTAR cathode up to 13.1A of discharge current and for the NEXIS cathode up to 30A of discharge current.

  4. Thrust Stand Measurements of a Conical Inductive Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Hallock, Ashley K.; Polzin, Kurt A.

    2013-01-01

    Inductive Pulsed Plasma Thrusters (iPPT) spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current Propellant is accelerated and expelled at a high exhaust velocity (O(10 -- 100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. A conical coil geometry may offer higher propellant utilization efficiency over that of a at inductive coil, however an increase in propellant utilization may be met with a decrease in axial electromagnetic acceleration, and in turn, a decrease in the total axially-directed kinetic energy imparted to the propellant.

  5. Design and Performance Estimates of an Ablative Gallium Electromagnetic Thruster

    NASA Technical Reports Server (NTRS)

    Thomas, Robert E.

    2012-01-01

    The present study details the high-power condensable propellant research being conducted at NASA Glenn Research Center. The gallium electromagnetic thruster is an ablative coaxial accelerator designed to operate at arc discharge currents in the range of 10-25 kA. The thruster is driven by a four-parallel line pulse forming network capable of producing a 250 microsec pulse with a 60 kA amplitude. A torsional-type thrust stand is used to measure the impulse of a coaxial GEM thruster. Tests are conducted in a vacuum chamber 1.5 m in diameter and 4.5 m long with a background pressure of 2 microtorr. Electromagnetic scaling calculations predict a thruster efficiency of 50% at a specific impulse of 2800 seconds.

  6. 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.

  7. 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.

  8. Magnetohydrodynamic MACH Code Used to Simulate Magnetoplasmadynamic Thrusters

    NASA Technical Reports Server (NTRS)

    Mikellides, Pavlos G.; LaPointe, Michael R.

    2002-01-01

    The On-Board Propulsion program at the NASA Glenn Research Center is utilizing a state of-the-art numerical simulation to model the performance of high-power electromagnetic plasma thrusters. Such thrusters are envisioned for use in lunar and Mars cargo transport, piloted interplanetary expeditions, and deep-space robotic exploration of the solar system. The experimental portion of this program is described in reference 1. This article describes the numerical modeling program used to guide the experimental research. The synergistic use of numerical simulations and experimental research has spurred the rapid advancement of high-power thruster technologies for a variety of bold new NASA missions. From its inception as a U.S. Department of Defense code in the mid-1980's, the Multiblock Arbitrary Coordinate Hydromagnetic (MACH) simulation tool has been used by the plasma physics community to model a diverse range of plasma problems--including plasma opening switches, inertial confinement fusion concepts, compact toroid formation and acceleration, z-pinch implosion physics, laser-target interactions, and a variety of plasma thrusters. The MACH2 code used at Glenn is a time-dependent, two-dimensional, axisymmetric, multimaterial code with a multiblock structure. MACH3, a more recent three-dimensional version of the code, is currently undergoing beta tests. The MACH computational mesh moves in an arbitrary Lagrangian-Eulerian (ALE) fashion that allows the simulation of diffusive-dominated and dispersive-dominated problems, and the mesh can be refined via a variety of adaptive schemes to capture regions of varying characteristic scale. The mass continuity and momentum equations model a compressible viscous fluid, and three energy equations are used to simulate nonthermal equilibrium between electrons, ions, and the radiation field. Magnetic fields are modeled by an induction equation that includes resistive diffusion, the Hall effect, and a thermal source for magnetic

  9. 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

  10. 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.

  11. 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

  12. 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.

  13. 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.

  14. 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

  15. High Power MPD Thruster Development at the NASA Glenn Research Center

    NASA Technical Reports Server (NTRS)

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

    2001-01-01

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

  16. 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.

  17. Rail accelerators for space transportation: An experimental investigation

    NASA Technical Reports Server (NTRS)

    Zana, L. M.; Kerslake, W. R.; Sturman, J. L.

    1986-01-01

    An experimental program was conducted at the Lewis Research Center with the objective of investigating the technical feasibility of rail accelerators for propulsion applications. Single-stage, plasma driven rail accelerators of small (4 by 6 mm) and medium (12.5 by 12.5 mm) bores were tested at peak accelerating currents of 50 to 450 kA. Streak-camera photography was used to provide a qualitative description of plasma armature acceleration. The effects of plasma blowby and varying bore pressure on the behavior of plasma armatures were studied.

  18. 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.

  19. Electrostatic Plasma Accelerator (EPA)

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Aston, Graeme

    1989-01-01

    The Electrostatic Plasma Accelerator (EPA) is a thruster concept which promises specific impulse levels between low power arcjets and those of the ion engine while retaining the relative simplicity of the arcjet. The EPA thruster produces thrust through the electrostatic acceleration of a moderately dense plasma. No accelerating electrodes are used and the specific impulse is a direct function of the applied discharge voltage and the propellant atomic mass. The goal of the present program is to demonstrate feasibility of the EPA thruster concept through experimental and theoretical investigations of the EPA acceleration mechanism and discharge chamber performance. Experimental investigations will include operating the test bed ion (TBI) engine as an EPA thruster and parametrically varying the thruster geometry and operating conditions to quantify the electrostatic plasma acceleration effect. The theoretical investigations will include the development of a discharge chamber model which describes the relationships between the engine size, plasma properties, and overall performance. For the EPA thruster to be a viable propulsion concept, overall thruster efficiencies approaching 30% with specific impulses approaching 1000 s must be achieved.

  20. Scaling laws for electromagnetic pulsed plasma thrusters

    NASA Astrophysics Data System (ADS)

    Ziemer, J. K.; Choueiri, E. Y.

    2001-08-01

    The scaling laws of pulsed plasma thrusters operating in the predominantly electromagnetic acceleration mode (EM-PPT) are investigated theoretically and experimentally using gas-fed pulsed plasma thrusters. A fundamental characteristic velocity that depends on the inductance per unit length and the square root of the capacitance to the initial inductance ratio is identified. An analytical model of the discharge current predicts scaling laws in which the propulsive efficiency is proportional to the EM-PPT performance scaling number, defined here as the ratio of the exhaust velocity to the EM-PPT characteristic velocity. The importance of the effective plasma resistance in improving the propulsive performance is shown. To test the validity of the predicted scaling relations, the performance of two gas-fed pulsed plasma thruster designs (one with coaxial electrodes and the other with parallel-plate electrodes), was measured under 70 different operating conditions using an argon plasma. The measurements demonstrate that the impulse bit scales linearly with the integral of the square of the discharge current as expected for an electromagnetic accelerator. The measured performance scaling is shown to be in good agreement with the theoretically predicted scaling. Normalizing the exhaust velocity and the impulse-to-energy ratio by the EM-PPT characteristic velocity collapses almost all the measured data onto single curves that uphold the general validity of these scaling laws. [12pt]This paper is dedicated to the memory of Dr Daniel Birx

  1. 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.

  2. 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.

  3. Throttling Impacts on Hall Thruster Performance, Erosion, and Qualification for NASA Science Missions

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.; DeHoyos, Amado

    2007-01-01

    With the SMART-1, Department of Defense, and commercial industry successes in Hall thruster technologies, NASA has started considering Hall thrusters for science missions. The recent Discovery proposals included a Hall thruster science mission and the In-Space Propulsion Project is investing in Hall thruster technologies. As the confidence in Hall thrusters improve, ambitious multi-thruster missions are being considered. Science missions often require large throttling ranges due to the 1/r(sup 2) power drop-off from the sun. Deep throttling of Hall thrusters will impact the overall system performance. Also, Hall thrusters can be throttled with both current and voltage, impacting erosion rates and performance. Last, electric propulsion thruster lifetime qualification has previously been conducted with long duration full power tests. Full power tests may not be appropriate for NASA science missions, and a combination of lifetime testing at various power levels with sufficient analysis is recommended. Analyses of various science missions and throttling schemes using the Aerojet BPT-4000 and NASA 103M HiVHAC thruster are presented.

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

    NASA Technical Reports Server (NTRS)

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

    2010-01-01

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

  5. 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.

  6. 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.

  7. Testing of the Qinetiq T6 Thruster in Support of the ESA BepiColombo Mercury Mission for the ESA BepiColombo Mission

    NASA Astrophysics Data System (ADS)

    Wallace, N.

    2004-10-01

    This paper presents the results of two test programmes. The first involved a series of simultaneous thruster firing tests performed at QinetiQ in the QinetiQ LEEP 2 (2nd Large European Electric Propulsion) vacuum test facility using 2 T6 thrusters operating between 50 and 230 mN at a specific impulse above 4500s. The test configuration allowed the simultaneous operation of 2 thrusters at a range of thrust levels and thruster spacings. The spacing between thrusters could be varied between 800 mm and 10 mm with the thrusters operating. One of the thrusters, designated the Primary, was suspended from a pendulum thrust balance and was aligned with a large diameter ion beam probe array. These diagnostics allowed the thrust to be directly measured and the ion beam plume to be characterised, allowing comparisons to be made between single and simultaneous operation at different thruster spacings. Testing was also performed with different combinations of neutralisers to determine the potential effects of dual thruster operation with a single neutraliser. The second series of tests involved the operation of a single T6 thruster at a thrust level of 208 mN at the elevated thermal conditions predicted for Mercury. Thermal and mechanical analyses of the T6 (including additional thrusters in close proximity) indicated the worst case thermal conditions occur when the thruster assembly was illuminated from the side, producing the most severe temperatures and thermal gradients across the thruster structure. A 500 hour test was performed in which thruster performance was assessed using electrical operating parameters, thrust balance and beam probe array measurements taken with the thruster operating at both nominal and simulated Mercury thermal conditions. The latter was produced using a metal collar equipped with electrical heaters, placed in close proximity to the thruster structure. The heater collar could be positioned and withdrawn without the need to vent the chamber and

  8. 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.

  9. Thrust Stand Measurements of a Conical Pulsed Inductive Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Hallock, Ashley K.; Polzin, Kurt A.; Emsellem, Gregory D.

    2012-01-01

    Pulsed inductive plasma thrusters [1-3] are spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current. Propellant is accelerated and expelled at a high exhaust velocity (O(10-100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, pulsed inductive plasma thrusters can su er from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. The Microwave Assisted Discharge Inductive Plasma Accelerator (MAD-IPA)[4], shown in Fig. 1 is a pulsed inductive plasma thruster that is able to operate at lower pulse energies by partially ionizing propellant with an electron cyclotron resonance (ECR) discharge inside a conical inductive coil whose geometry serves to potentially increase propellant and plasma plume containment relative to at coil geometries. The ECR plasma is created with the use of permanent mag- nets arranged to produce a thin resonance region along the inner surface of the coil, restricting plasma formation and, in turn, current sheet formation to areas of high magnetic coupling to the driving coil.

  10. Design and Testing of a Small Inductive Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Martin, Adam K.; Dominguez, Alexandra; Eskridge, Richard H.; Polzin, Kurt A.; Riley, Daniel P.; Perdue, Kevin A.

    2015-01-01

    The design and testing of a small inductive pulsed plasma thruster (IPPT) is described. The device was built as a test-bed for the pulsed gas-valves and solid-state switches required for a thruster of this kind, and was designed to be modular to facilitate modification. The thruster in its present configuration consists of a multi-turn, spiral-wound acceleration coil (270 millimeters outer diameter, 100 millimeters inner diameter) driven by a 10 microfarad capacitor and switched with a high-voltage thyristor, a propellant delivery system including a fast pulsed gas-valve, and a glow-discharge pre-ionizer circuit. The acceleration coil circuit may be operated at voltages up to 4 kilovolts (the thyristor limit is 4.5 kilovolts) and the thruster operated at cyclic-rates up to 30 Herz. Initial testing of the thruster, both bench-top and in-vacuum, has been performed. Cyclic operation of the complete device was demonstrated (at 2 Herz), and a number of valuable insights pertaining to the design of these devices have been gained.

  11. 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.

  12. 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

  13. 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.

  14. Numerical Simulation of Ion Thruster Plume Backflow for Spacecraft Contamination Assessment.

    NASA Astrophysics Data System (ADS)

    Samanta Roy, Robie I.

    1995-01-01

    A study is presented of the induced environment produced by an ion thruster, and its interactions with spacecraft. Axisymmetric and fully three-dimensional physical and numerical models of an ion thruster plume are developed and utilized to understand the physical processes involved, as well as to make useful predictions of spacecraft contamination. Components included in the model are primary beam ions, neutral propellant efflux, slow propellant ions created mainly by charge-exchange (CEX) collisions, non-propellant efflux (NPE) such as sputtered molybdenum grid metal, and neutralizing electrons. Primary focus is on the creation and transport of both the CEX propellant ions created from the beam ions and neutral propellant, and charged NPE from the thruster plume into the backflow region. The numerical model utilizes the hybrid plasma particle-in-cell (PIC) method, and the fully three-dimensional implementation is designed for multi -computer environments. Computational results of the CEX ion density and flow angles show good qualitative and quantitative agreement with experimental data. It is shown that the CEX ions created in the beam accelerate outwards and form two distinct energy populations, one with a significant backstreaming component. The effect of the background tank pressure in ground experiments, and the accelerator grid impingement current is examined. The backflow contamination from NASA's current 30 cm xenon ion thruster is investigated over the operating envelope of the thruster, and predictions for space operation are made. Scaling relationships for the backflow previously identified are confirmed. Issues regarding the electron temperature in the plume are explored, and it is shown that backflow contamination increases with electron temperature. Fully three-dimensional results with up to 17.5 million particles are presented. It is shown that the spacecraft geometry plays a strong role in the expansion of the CEX plasma. The contamination from the

  15. 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.

  16. 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.

  17. Recent Progress in Silicon-Based MEMS Field Emission Thrusters

    NASA Astrophysics Data System (ADS)

    Lenard, Roger X.; Kravitz, Stanley H.; Tajmar, Martin

    2005-02-01

    The Indium Field Emission Thruster (In-FET) is a highly characterized and space-proven device based on space-qualified liquid metal ion sources. There is also extensive experience with liquid metal ion sources for high-brightness semiconductor fabrications and inspection Like gridded ion engines, In-FETs efficiently accelerate ions through a series of high voltage electrodes. Instead of a plasma discharge to generate ions, which generates a mixture of singly and doubly charged ions as well as neutrals, indium metal is melted (157°C) and fed to the tip of a capillary tube where very high local electric fields perform more-efficient field emission ionization, providing nearly 100% singly charged species. In-FETs do not have the associated losses or lifetime concerns of a magnetically confined discharge and hollow cathode in ion thrusters. For In-FETs, propellant efficiencies ˜100% stipulate single-emitter currents ⩽10μA, perhaps as low as 5μA of current. This low emitter current results in ⩽0.5 W/emitter. Consequently, if the In-FET is to be used for future Human and Robotic missions under President Bush's Exploration plan, a mechanism to generate very high power levels is necessary. Efficient high-power operation requires many emitter/extractor pairs. Conventional fabrication techniques allow 1-10 emitters in a single module, with pain-staking precision required. Properly designed and fabricated In-FETs possess electric-to-jet efficiency >90% and a specific mass <0.25 kg/kWe. MEMS techniques allow reliable batch processing with ˜160,000 emitters in a 10×10-cm array. Developing a 1.5kW 10×10-cm module is a necessary stepping-stone for >500 kWe systems where groups of 9 or 16 modules, with a single PPU/feed system, form the building blocks for even higher-power exploration systems. In 2003, SNL and ARCS produced a MEMS-based In-FET 5×5 emitter module with individually addressable emitter/extractor pairs on a 15×15mm wafer. The first MEMS thruster

  18. Colloid micro-Newton thruster development for the ST7-DRS and LISA missions

    NASA Technical Reports Server (NTRS)

    Ziemer, John K.; Gamero-Castano, Manuel; Hruby, Vlad; Spence, Doug; Demmons, Nate; McCormick, Ryan; Roy, Tom

    2005-01-01

    We present recent progress and development of the Busek Colloid Micro-Newton Thruster (CMNT) for the Space Technology 7 Disturbance Reduction System (ST7-DRS) and Laser Interferometer Space Antenna (LISA) Missions.

  19. Ram accelerator direct launch system for space cargo

    NASA Technical Reports Server (NTRS)

    Bruckner, A. P.; Hertzberg, A.

    1987-01-01

    The ram accelerator, a chemically-propelled mass driver, is presented as a new approach for directly launching acceleration-insensitive pay-loads into LEO. The cargo vehicle resembles the centerbody of a conventional ramjet and travels through a launch tube filled with a premixed gaseous fuel and oxidizer mixture. The tube acts as the outer cowling of the ramjet and the combustion process travels with the vehicle. Two modes of ram accelerator drive are described, which when used in sequence, are capable of accelerating the cargo vehicle to 10 km/sec. The requirements for placing a 2000 kg vehicle with 50 percent payload fraction into a 400 km orbit, with a minimum of on-board rocket propellant for circularization maneuvers, are examined. It is shown that aerodynamic heating during atmospheric transit results in very little ablation of the nose. Both direct and indirect orbital insertion scenarios are investigated, and a three-step maneuver consisting of two burns and aerobraking is found to minimize the on-board propellant mass. A scenario involving a parking orbit below the desired final orbit is suggested as a means to increase the flexibility of the mass launch concept.

  20. Graviresponses in Paramecium biaurelia under different accelerations: studies on the ground and in space.

    PubMed

    Hemmersbach, R; Voormanns, R; Hader, D P

    1996-10-01

    Behavioural responses to different accelerations below 1 g and up to 5 g were investigated in Paramecium biaurelia by using a centrifuge microscope on Earth and in space during a recent space flight. Increased stimulation (hypergravity) enhanced the negative gravitactic and the gravikinetic responses in Paramecium biaurelia within seconds. Cells did not adapt to altered gravitational conditions. Repetitive stimulation did not change the graviresponses. The minimum acceleration found to induce gravitaxis was between 0.16 and 0.3 g.

  1. ECR-GDM Thruster for Fusion Propulsion

    SciTech Connect

    Brainerd, Jerome J.; Reisz, Al

    2009-03-16

    The concept of the Gasdynamic Mirror (GDM) device for fusion propulsion was proposed by and Lee (1995) over a decade ago and several theoretical papers has supported the feasibility of the concept. A new ECR plasma source has been built to supply power to the GDM experimental thruster previously tested at the Marshall Space Flight Center (MSFC). The new plasma generator, powered by microwaves at 2.45 or 10 GHz. is currently being tested. This ECR plasma source operates in a number of distinct plasma modes, depending upon the strength and shape of the local magnetic field. Of particular interest is the compact plasma jet issuing form the plasma generator when operated in a mirror configuration. The measured velocity profile in the jet plume is bimodal, possibly as a result of the GDM effect in the ECR chamber of the thruster.

  2. Energy storage systems for MPD thrusters

    NASA Technical Reports Server (NTRS)

    Gabriel, S. B.

    1981-01-01

    Because of its high thrust density, the magnetoplasmadynamic (MPD) thruster is a promising candidate for many advanced space missions. Its high power requirements lead to operation in a pulsed mode using an intermediate energy storage device. The characteristics of a system consisting of a solar array, energy storage capacitor, and MPD thruster are studied for array powers in the range 25-375 kw. Assuming simple analytic models for the circuit components, the circuit charge and discharge equations are solved numerically, resulting in the system efficiency and capacitance. The system efficiency is inversely proportional to array power and decreases with circuit resistance. Alternative methods of energy storage such as a pulse forming network and a homopolar generator, are presented, and an overall comparison between all of the methods is given.

  3. Advanced Hall Electric Propulsion for Future In-space Transportation

    NASA Technical Reports Server (NTRS)

    Oleson, Steven R.; Sankovic, John M.

    2001-01-01

    The Hall thruster is an electric propulsion device used for multiple in-space applications including orbit raising, on-orbit maneuvers, and de-orbit functions. These in-space propulsion functions are currently performed by toxic hydrazine monopropellant or hydrazine derivative/nitrogen tetroxide bi-propellant thrusters. The Hall thruster operates nominally in the 1500 sec specific impulse regime. It provides greater thrust to power than conventional gridded ion engines, thus reducing trip times and operational life when compared to that technology in Earth orbit applications. The technology in the far term, by adding a second acceleration stage, has shown promise of providing over 4000s Isp, the regime of the gridded ion engine and necessary for deep space applications. The Hall thruster system consists of three parts, the thruster, the power processor, and the propellant system. The technology is operational and commercially available at the 1.5 kW power level and 5 kW application is underway. NASA is looking toward 10 kW and eventually 50 kW-class engines for ambitious space transportation applications. The former allows launch vehicle step-down for GEO missions and demanding planetary missions such as Europa Lander, while the latter allows quick all-electric propulsion LEO to GEO transfers and non-nuclear transportation human Mars missions.

  4. 4.5-kW Hall Effect Thruster Evaluated

    NASA Technical Reports Server (NTRS)

    Mason, Lee S.

    2000-01-01

    As part of an Interagency Agreement with the Air Force Research Lab (AFRL), a space simulation test of a Russian SPT 140 Hall Effect Thruster was completed in September 1999 at Vacuum Facility 6 at the NASA Glenn Research Center at Lewis Field. The thruster was subjected to a three-part test sequence that included thrust and performance characterization, electromagnetic interference, and plume contamination. SPT 140 is a 4.5-kW thruster developed under a joint agreement between AFRL, Atlantic Research Corp, and Space Systems/Loral, and was manufactured by the Fakal Experimental Design Bureau of Russia. All objectives were satisfied, and the thruster performed exceptionally well during the 120-hr test program, which comprised 33 engine firings. The Glenn testing provided a critical contribution to the thruster development effort, and the large volume and high pumping speed of this vacuum facility was key to the test s success. The low background pressure (1 10 6 torr) provided a more accurate representation of space vacuum than is possible in most vacuum chambers. The facility had been upgraded recently with new cryogenic pumps and sputter shielding to support the active electric propulsion program at Glenn. The Glenn test team was responsible for all test support equipment, including the thrust stand, power supplies, data acquisition, electromagnetic interference measurement equipment, and the contamination measurement system.

  5. Particle-in-cell simulations of Hall plasma thrusters

    NASA Astrophysics Data System (ADS)

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

    2016-07-01

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

  6. 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

  7. Segmented electrode hall thruster with reduced plume

    DOEpatents

    Fisch, Nathaniel J.; Raitses, Yevgeny

    2004-08-17

    An apparatus and method for thrusting plasma, utilizing a Hall thruster with segmented electrodes along the channel, which make the acceleration region as localized as possible. Also disclosed are methods of arranging the electrodes so as to minimize erosion and arcing. Also disclosed are methods of arranging the electrodes so as to produce a substantial reduction in plume divergence. The use of electrodes made of emissive material will reduce the radial potential drop within the channel, further decreasing the plume divergence. Also disclosed is a method of arranging and powering these electrodes so as to provide variable mode operation.

  8. 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.

  9. Effect of Background Pressure on the Plasma Oscillation Characteristics of the HiVHAc Hall Thruster

    NASA Technical Reports Server (NTRS)

    Huang, Wensheng; Kamhawi, Hani; Lobbia, Robert B.; Brown, Daniel L.

    2014-01-01

    During a component compatibility test of the NASA HiVHAc Hall thruster, a number of plasma diagnostics were implemented to study the effect of varying facility background pressure on thruster operation. These diagnostics characterized the thruster performance, the plume, and the plasma oscillations in the thruster. Thruster performance and plume characteristics as functions of background pressure were previously published. This paper focuses on changes in the plasma oscillation characteristics with changing background pressure. The diagnostics used to study plasma oscillations include a high-speed camera and a set of high-speed Langmuir probes. The results show a rise in the oscillation frequency of the "breathing" mode with rising background pressure, which is hypothesized to be due to a shortening acceleration/ionization zone. An attempt is made to apply a simplified ingestion model to the data. The combined results are used to estimate the maximum acceptable background pressure for performance and wear testing.

  10. Measurement and Data Distribution for Microgravity Accelerations on the International Space Station

    NASA Technical Reports Server (NTRS)

    McPherson, Kevin; Hrovat, Kenneth

    1999-01-01

    Two accelerometer systems will be available on the International Space Station to support microgravity payloads with information about the quasi-steady and vibratory acceleration environment of the research facilities. The Microgravity Acceleration Measurement System will record contributions to the quasi-steady microgravity environment, including the influences of aerodynamic drag, vehicle rotation, and venting effects. The Space Acceleration Measurement System-II will measure vibratory disturbances on-board due to vehicle, crew, and equipment disturbances. Due to the dynamic nature of the microgravity environment and its potential to influence sensitive experiments, NASA's Principal Investigator Microgravity Services project has initiated a plan through which the data from these instruments will be distributed to researchers in a timely and meaningful fashion. Beyond the obvious benefit of correlation between accelerations and the scientific phenomena being studied, such information is also useful for hardware developers who can gain qualitative and quantitative feedback about their facility acceleration output to station.

  11. Preliminary tests of the electrostatic plasma accelerator

    NASA Technical Reports Server (NTRS)

    Aston, G.; Acker, T.

    1990-01-01

    This report describes the results of a program to verify an electrostatic plasma acceleration concept and to identify those parameters most important in optimizing an Electrostatic Plasma Accelerator (EPA) thruster based upon this thrust mechanism. Preliminary performance measurements of thrust, specific impulse and efficiency were obtained using a unique plasma exhaust momentum probe. Reliable EPA thruster operation was achieved using one power supply.

  12. A Very-High-Specific-Impulse Relativistic Laser Thruster

    SciTech Connect

    Horisawa, Hideyuki; Kimura, Itsuro

    2008-04-28

    Characteristics of compact laser plasma accelerators utilizing high-power laser and thin-target interaction were reviewed as a potential candidate of future spacecraft thrusters capable of generating relativistic plasma beams for interstellar missions. Based on the special theory of relativity, motion of the relativistic plasma beam exhausted from the thruster was formulated. Relationships of thrust, specific impulse, input power and momentum coupling coefficient for the relativistic plasma thruster were derived. It was shown that under relativistic conditions, the thrust could be extremely large even with a small amount of propellant flow rate. Moreover, it was shown that for a given value of input power thrust tended to approach the value of the photon rocket under the relativistic conditions regardless of the propellant flow rate.

  13. Correlation of ion and beam current densities in Kaufman thrusters.

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.

    1973-01-01

    In the absence of direct impingement erosion, electrostatic thruster accelerator grid lifetime is defined by the charge exchange erosion that occurs at peak values of the ion beam current density. In order to maximize the thrust from an engine with a specified grid lifetime, the ion beam current density profile should therefore be as flat as possible. Knauer (1970) has suggested this can be achieved by establishing a radial plasma uniformity within the thruster discharge chamber; his tests with the radial field thruster provide an example of uniform plasma properties within the chamber and a flat ion beam profile occurring together. It is shown that, in particular, the ion density profile within the chamber determines the beam current density profile, and that a uniform ion density profile at the screen grid end of the discharge chamber should lead to a flat beam current density profile.

  14. Scaling relationships for mercury and gaseous propellant ion thrusters

    NASA Technical Reports Server (NTRS)

    Wilbur, P. J.; Kaufman, H. R.

    1978-01-01

    A model describing the doubly charged ion densities in argon and xenon ion thrusters is presented. Doubly charged ions are shown to be produced in significant numbers from both the neutral and singly ionized states. Doubly-to-singly charged ion density ratios calculated using this model are compared to experimental values of this ratio measured in the 15 cm multipole thruster using both propellants. Agreement between theory and experiment is shown to be good. Using this doubly charged ion model, and a similar one obtained previously for mercury, together with the constant neutral loss rate theory, Child's current density law and accelerator grid fabrication limitations, a set of scaling relationships are developed. These relationships show the propellant utilizations, thrust densities and discharge power levels that can be expected as thruster diameter and/or specific impulse are increased and doubly charged ion densities are held at acceptably low values.

  15. 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.

  16. Deformed phase space Kaluza-Klein cosmology and late time acceleration

    NASA Astrophysics Data System (ADS)

    Sabido, M.; Yee-Romero, C.

    2016-06-01

    The effects of phase space deformations on Kaluza-Klein cosmology are studied. The deformation is introduced by modifying the symplectic structure of the minisuperspace variables. In the deformed model, we find an accelerating scale factor and therefore infer the existence of an effective cosmological constant from the phase space deformation parameter β.

  17. 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.

  18. The Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

    Martin, Adam; Eskridge, Richard; Fimognan, Peter; Koelfgen, Syri J.; Lee, Mike

    2004-01-01

    A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic fields. If B(sup p)/B(sub t) is much greater than 1, it is an FRC; if B(sub p) approximately equals B(sub t), it is a Spheromak. A plasmoid thruster would operate by repetitively producing plasmoids that are accelerated to high velocity. The process is inductive, and the magnetic structure of the plasmoid suppresses thermal and mass losses, and improves detachment of the exhaust. This concept should be capable of producing an Isp in the range of 5,000 - 10,OOO seconds, with high thrust density. PTX is a device designed to study this concept. The plasmoid is formed inside of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank. Experiments conducted with a static-fill of propellant gas (6% H2 in He) demonstrated reliable ionization over a pressure range of 40 - 200 mTorr. A fast gas-puff valve to inject propellant has since been added, and a ringing pre-ionization circuit to independently control ionization has been tested. Hydrogen, deuterium, argon, and an N2/H2 mixture have been tried as propellants. Measurements of the plasmoid shape, mass, and velocity, using a variety of diagnostics will be presented,

  19. The Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

    Martin, Adam; Eskridge, Richard; Lee, Mike; Fimohnsti, Peter; Koelfgen, Syri J.

    2005-01-01

    A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic fields. If B(sub p/B(sub t) much greater than 1 it is an FRC; if B(sub p) approximately equal to B(sub t), it is a Spheromak. A plasmoid thruster would operate by repetitively producing plasmoids that are accelerated to high velocity. The process is inductive, and the magnetic structure of the plasmoid suppresses thermal and mass losses, and improves detachment of the exhaust, This concept should be capable of producing an Isp in the range of 5,000 - l0,000 seconds, with high thrust density. PTX is a device designed to study this concept. The plasmoid is formed inside of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank. Experiments conducted with a static-fill of propellant gas (6% H2 in He) demonstrated liable ionization over a pressure range of 40 - 200 mTorr. A fast gas-puff valve to inject propellant has since been added, and a ringing preionization circuit to independently control ionization has been tested, hydrogen, deuterium, argon, and an N2 / H2 mixture have been tried as propellants. Measurements of the plasmoid shape, mass, and velocity, using a variety of diagnostics will be presented.

  20. In-Space Propulsion Solar Electric Propulsion Technology Overview

    NASA Astrophysics Data System (ADS)

    Dankanich, John W.

    2006-12-01

    NASA’s In-space Propulsion Technology Project is developing new propulsion technologies that can enable or enhance near and mid-term NASA science missions. The solar electric propulsion technology area has been investing in NASA’s Evolutionary Xenon Thruster (NEXT), the High Voltage Hall Accelerator (HiVHAC), lightweight reliable feed systems, wear testing and thruster modeling. These investments are specifically targeted to increase planetary science payload capability, expand the envelope of planetary science destinations, and significantly reduce the travel times, risk and cost of NASA planetary science missions. Current status and expected capabilities of the solar electric propulsion technologies will be discussed.

  1. Inductive Pulsed Plasma Thruster Development and Testing at NASA-MSFC

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.

    2013-01-01

    THE inductive pulsed plasma thruster (IPPT) is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. In the present work, we present a summary of the IPPT research and development conducted at NASA's Marshall Space Flight Center (MSFC). As a higher-power, still relatively low readiness level system, there are many issues associated with the eventual deployment and use of the IPPT as a primary propulsion system on spacecraft that remain to be addressed. The present program aimed to fabricate and test hardware to explore how these issues could be addressed. The following specific areas were addressed within the program and will be discussed within this paper. a) Conical theta-pinch IPPT geometry thruster configuration. b) Repetition-rate multi-kW thruster pulsing. c) Long-lifetime pulsed gas valve. d) Fast pulsed gas valve driver and controller. e) High-voltage, repetitive capacitor charging power processing unit. During the course of testing, a number of specific tests were conducted, including several that, to our knowledge, have either never been previously conducted (such as multi-KW repetition-rate operation) or have not been performed since the early 1990s (direct IPPT thrust measurements).2 Conical theta-pinch IPPT thrust stand measurements are presented in Fig. 1 while various time-integrated and time

  2. ACCELERATORS: Matching by solenoids in space charge dominated LEBTs

    NASA Astrophysics Data System (ADS)

    Li, Jin-Hai; Tang, Jing-Yu; Ouyang, Hua-Fu

    2009-10-01

    The betatron matching of a rotationally asymmetric beam in space charge dominated low-energy beam transports (LEBTs) where solenoids are used for the transverse matching has been studied. For better understanding, the coupling elements of a beam matrix are interpreted in special forms that are products of a term defined by the Larmor rotation angle and another by the difference between the beam matrix elements in the two transverse planes. The coupling form originally derived from the rotationally symmetric field in solenoids still holds when taking into account the rotationally asymmetric space charge forces that are due to the unequal emittance in the two transverse planes. It is shown in this paper that when an LEBT mainly comprising solenoids transports a beam having unequal emittance in the two transverse planes and the linear space charge force is taken into account, the initial Twiss parameters can be modified to obtain the minimum and equal emittance at the LEBT exit. The TRACE3D calculations also prove the principle. However, when quadrupoles that are also rotationally asymmetric are involved in between solenoids, the coupling between the two transverse planes becomes more complicated and the emittance increase is usually unavoidable. A matching example using the CSNS (China Spallation Neutron Source) LEBT conditions is also presented.

  3. Accelerating molecular property calculations with nonorthonormal Krylov space methods

    NASA Astrophysics Data System (ADS)

    Furche, Filipp; Krull, Brandon T.; Nguyen, Brian D.; Kwon, Jake

    2016-05-01

    We formulate Krylov space methods for large eigenvalue problems and linear equation systems that take advantage of decreasing residual norms to reduce the cost of matrix-vector multiplication. The residuals are used as subspace basis without prior orthonormalization, which leads to generalized eigenvalue problems or linear equation systems on the Krylov space. These nonorthonormal Krylov space (nKs) algorithms are favorable for large matrices with irregular sparsity patterns whose elements are computed on the fly, because fewer operations are necessary as the residual norm decreases as compared to the conventional method, while errors in the desired eigenpairs and solution vectors remain small. We consider real symmetric and symplectic eigenvalue problems as well as linear equation systems and Sylvester equations as they appear in configuration interaction and response theory. The nKs method can be implemented in existing electronic structure codes with minor modifications and yields speed-ups of 1.2-1.8 in typical time-dependent Hartree-Fock and density functional applications without accuracy loss. The algorithm can compute entire linear subspaces simultaneously which benefits electronic spectra and force constant calculations requiring many eigenpairs or solution vectors. The nKs approach is related to difference density methods in electronic ground state calculations and particularly efficient for integral direct computations of exchange-type contractions. By combination with resolution-of-the-identity methods for Coulomb contractions, three- to fivefold speed-ups of hybrid time-dependent density functional excited state and response calculations are achieved.

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

    SciTech Connect

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

    2012-12-15

    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. 15 cm cusped magnetic field mercury ion thruster research

    NASA Technical Reports Server (NTRS)

    Beattie, J. R.; Wilbur, P. J.

    1975-01-01

    The importance of achieving a uniform current density in the exhaust beam of an electrostatic ion thruster is discussed in terms of thrust level and accelerator grid lifetime. A neutral residence time approach is used to propose a magnetic field geometry which should produce a highly uniform beam current density. The discharge chamber length to diameter ratio is shown to be an important optimization parameter and experimental evaluation of the cusped field thruster over a wide range of this parameter is presented. Beam profile measurements 5 cm downstream of the accelerator grid indicate a beam profile flatness parameter which is 25% greater than the SERT II value. Flatness parameters extrapolated to the plane of the accelerator grid are demonstrated to be as high as 0.9.

  6. 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.

  7. Accelerating molecular property calculations with nonorthonormal Krylov space methods

    DOE PAGESBeta

    Furche, Filipp; Krull, Brandon T.; Nguyen, Brian D.; Kwon, Jake

    2016-05-03

    Here, we formulate Krylov space methods for large eigenvalue problems and linear equation systems that take advantage of decreasing residual norms to reduce the cost of matrix-vector multiplication. The residuals are used as subspace basis without prior orthonormalization, which leads to generalized eigenvalue problems or linear equation systems on the Krylov space. These nonorthonormal Krylov space (nKs) algorithms are favorable for large matrices with irregular sparsity patterns whose elements are computed on the fly, because fewer operations are necessary as the residual norm decreases as compared to the conventional method, while errors in the desired eigenpairs and solution vectors remainmore » small. We consider real symmetric and symplectic eigenvalue problems as well as linear equation systems and Sylvester equations as they appear in configuration interaction and response theory. The nKs method can be implemented in existing electronic structure codes with minor modifications and yields speed-ups of 1.2-1.8 in typical time-dependent Hartree-Fock and density functional applications without accuracy loss. The algorithm can compute entire linear subspaces simultaneously which benefits electronic spectra and force constant calculations requiring many eigenpairs or solution vectors. The nKs approach is related to difference density methods in electronic ground state calculations, and particularly efficient for integral direct computations of exchange-type contractions. By combination with resolution-of-the-identity methods for Coulomb contractions, three- to fivefold speed-ups of hybrid time-dependent density functional excited state and response calculations are achieved.« less

  8. The Simpsons program 6-D phase space tracking with acceleration

    SciTech Connect

    Machida, S. )

    1993-12-25

    A particle tracking code, Simpsons, in 6-D phase space including energy ramping has been developed to model proton synchrotrons and storage rings. We take time as the independent variable to change machine parameters and diagnose beam quality in a quite similar way as real machines, unlike existing tracking codes for synchrotrons which advance a particle element by element. Arbitrary energy ramping and rf voltage curves as a function of time are read as an input file for defining a machine cycle. The code is used to study beam dynamics with time dependent parameters. Some of the examples from simulations of the Superconducting Super Collider (SSC) boosters are shown.

  9. The Simpsons program 6-D phase space tracking with acceleration

    SciTech Connect

    Machida, S.

    1993-02-01

    A particle tracking code, Simpsons, in 6-D phase space including energy ramping has been developed to model proton synchrotrons and storage rings. We take time as the independent variable to change machine parameters and diagnose beam quality in a quite similar way as real machines, unlike existing tracking codes for synchrotrons which advance a particle element by element. Arbitrary energy ramping and rf voltage curves as a function of time are read as an input file for defining a machine cycle. The code is used to study beam dynamics with time dependent parameters. Some of the examples from simulations of the Superconducting Super Collider (SSC) boosters are shown.

  10. The Simpsons program 6-D phase space tracking with acceleration

    NASA Astrophysics Data System (ADS)

    Machida, S.

    1993-12-01

    A particle tracking code, Simpsons, in 6-D phase space including energy ramping has been developed to model proton synchrotrons and storage rings. We take time as the independent variable to change machine parameters and diagnose beam quality in a quite similar way as real machines, unlike existing tracking codes for synchrotrons which advance a particle element by element. Arbitrary energy ramping and rf voltage curves as a function of time are read as an input file for defining a machine cycle. The code is used to study beam dynamics with time dependent parameters. Some of the examples from simulations of the Superconducting Super Collider (SSC) boosters are shown.

  11. Reasonable structure for the discharge type plasma source. [In Pulsed Plasma Thruster

    SciTech Connect

    An, S.M.

    1987-01-01

    Experiments conducted with a magnetoplasma thruster in which plasma production and acceleration were treated separately indicate that different plasma source geometries have the most direct effect on energy conversion efficiency. An analysis of cup and tube type constraining structures shows the cup type to incur the greatest losses. It is noted that a parallel rail-type open structure such as that employed by the Chinese MDT-2A thruster leads to substantial discharge process dispersion. It is emphasized that the type and performance characteristics of a plasma source have a critical influence on thruster behavior. 5 references.

  12. 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.

  13. Cosmic-Ray Accelerators in Milky Way studied with the Fermi Gamma-ray Space Telescope

    SciTech Connect

    Kamae, Tuneyoshi; /SLAC /KIPAC, Menlo Park

    2012-05-04

    High-energy gamma-ray astrophysics is now situated at a confluence of particle physics, plasma physics and traditional astrophysics. Fermi Gamma-ray Space Telescope (FGST) and upgraded Imaging Atmospheric Cherenkov Telescopes (IACTs) have been invigorating this interdisciplinary area of research. Among many new developments, I focus on two types of cosmic accelerators in the Milky-Way galaxy (pulsar, pulsar wind nebula, and supernova remnants) and explain discoveries related to cosmic-ray acceleration.

  14. 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.

  15. Thrust Stand Measurements of a Conical Inductive Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Hallock, Ashley K.; Polzin, Kurt A.

    2012-01-01

    Inductive Pulsed Plasma Thrusters (iPPT) are spacecraft propulsion devices in which electrical energy is capacitively stored and then discharged through an inductive coil. The thruster is electrodeless, with a time-varying current in the coil interacting with a plasma covering the face of the coil to induce a plasma current. Propellant is accelerated and expelled at a high exhaust velocity (O(10 .. 100 km/s)) by the Lorentz body force arising from the interaction of the magnetic field and the induced plasma current. While this class of thruster mitigates the life-limiting issues associated with electrode erosion, inductive pulsed plasma thrusters can suffer from both high pulse energy requirements imposed by the voltage demands of inductive propellant ionization, and low propellant utilization efficiencies. A conical coil geometry may o er higher propellant utilization efficiency over that of a at inductive coil, however an increase in propellant utilization may be met with a decrease in axial electromagnetic acceleration, and in turn, a decrease in the total axially-directed kinetic energy imparted to the propellant.

  16. Correlated histogram representation of Monte Carlo derived medical accelerator photon-output phase space

    DOEpatents

    Schach Von Wittenau, Alexis E.

    2003-01-01

    A method is provided to represent the calculated phase space of photons emanating from medical accelerators used in photon teletherapy. The method reproduces the energy distributions and trajectories of the photons originating in the bremsstrahlung target and of photons scattered by components within the accelerator head. The method reproduces the energy and directional information from sources up to several centimeters in radial extent, so it is expected to generalize well to accelerators made by different manufacturers. The method is computationally both fast and efficient overall sampling efficiency of 80% or higher for most field sizes. The computational cost is independent of the number of beams used in the treatment plan.

  17. Accelerated space object tracking via graphic processing unit

    NASA Astrophysics Data System (ADS)

    Jia, Bin; Liu, Kui; Pham, Khanh; Blasch, Erik; Chen, Genshe

    2016-05-01

    In this paper, a hybrid Monte Carlo Gauss mixture Kalman filter is proposed for the continuous orbit estimation problem. Specifically, the graphic processing unit (GPU) aided Monte Carlo method is used to propagate the uncertainty of the estimation when the observation is not available and the Gauss mixture Kalman filter is used to update the estimation when the observation sequences are available. A typical space object tracking problem using the ground radar is used to test the performance of the proposed algorithm. The performance of the proposed algorithm is compared with the popular cubature Kalman filter (CKF). The simulation results show that the ordinary CKF diverges in 5 observation periods. In contrast, the proposed hybrid Monte Carlo Gauss mixture Kalman filter achieves satisfactory performance in all observation periods. In addition, by using the GPU, the computational time is over 100 times less than that using the conventional central processing unit (CPU).

  18. 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.

  19. 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.

  20. Evaluation of externally heated pulsed MPD thruster cathodes

    NASA Technical Reports Server (NTRS)

    Myers, Roger M.; Domonkos, Matthew; Gallimore, Alec D.

    1993-01-01

    Recent interest in solar electric orbit transfer vehicles (SEOTV's) has prompted a reevaluation of pulsed magnetoplasmadynamic (MPD) thruster systems due to their ease of power scaling and reduced test facility requirements. In this work the use of externally heated cathodes was examined in order to extend the lifetime of these thrusters to the 1000 to 3000 hours required for SEOTV missions. A pulsed MPD thruster test facility was assembled, including a pulse-forming network (PFN), ignitor supply and propellant feed system. Results of cold cathode tests used to validate the facility, PFN, and propellant feed system design are presented, as well as a preliminary evaluation of externally heated impregnated tungsten cathodes. The cold cathode thruster was operated on both argon and nitrogen propellants at peak discharge power levels up to 300 kW. The results confirmed proper operation of the pulsed thruster test facility, and indicated that large amounts of gas were evolved from the BaO-CaO-Al2O3 cathodes during activation. Comparison of the expected space charge limited current with the measured vacuum current when using the heated cathode indicate that either that a large temperature difference existed between the heater and the cathode or that the surface work function was higher than expected.

  1. Plasma Sheet Velocity Measurement Techniques for the Pulsed Plasma Thruster SIMP-LEX

    NASA Technical Reports Server (NTRS)

    Nawaz, Anuscheh; Lau, Matthew

    2011-01-01

    The velocity of the first plasma sheet was determined between the electrodes of a pulsed plasma thruster using three measurement techniques: time of flight probe, high speed camera and magnetic field probe. Further, for time of flight probe and magnetic field probe, it was possible to determine the velocity distribution along the electrodes, as the plasma sheet is accelerated. The results from all three techniques are shown, and are compared for one thruster geometry.

  2. Progress on the PT-1 Prototype Plasmoid Thruster

    NASA Technical Reports Server (NTRS)

    Eskridge, Richard H.; Martin, Adam K.

    2007-01-01

    The design and construction of a plasmoid thruster prototype is described. This thruster operates by expelling inductively formed plasmoids at high velocities. These plasmoids are field reversed configuration plasmas which are formed by reversing a magnetic flux frozen in an ionized gas inside a theta-pinch coil. The pinch coil is a unique multi-turn, multi-lead design chosen for optimization of inductance and field uniformity. A table-top bread-board demonstrator has been built at MSFC, and will be delivered to Radiance Technologies Inc. for further testing at the Auburn Space Power Institute.

  3. 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.

  4. 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.

  5. Thruster momentum transfer studies and magnetoplasmadynamic (MPD) thruster use in materials/surface modification

    NASA Technical Reports Server (NTRS)

    Thompson, E.; Collins, W. E.; Burger, A.; George, M. A.; Conner, J. D.

    1995-01-01

    This research project involves the systematic study of the MPD thruster for dual uses. Though it was designed as a thruster for space vehicles, the characteristics of the plasma make it an excellent candidate for industrial applications. This project will seek to characterize the system for use in materials processing and characterization. Crystals grown at Fisk University for use with solid state detectors will be studied. Surface modification on ZnCdTe, CdTe, and ZnTe will be studied using AFM, XPS and SAES. In addition to the surface modification studies, design work on a momentum transfer measurement device has been completed. The design and limitations of the device will be presented.

  6. 2-D Magnetohydrodynamic Modeling of A Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Thio, Y. C. Francis; Cassibry, J. T.; Wu, S. T.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    Experiments are being performed on the NASA Marshall Space Flight Center (MSFC) MK-1 pulsed plasma thruster. Data produced from the experiments provide an opportunity to further understand the plasma dynamics in these thrusters via detailed computational modeling. The detailed and accurate understanding of the plasma dynamics in these devices holds the key towards extending their capabilities in a number of applications, including their applications as high power (greater than 1 MW) thrusters, and their use for producing high-velocity, uniform plasma jets for experimental purposes. For this study, the 2-D MHD modeling code, MACH2, is used to provide detailed interpretation of the experimental data. At the same time, a 0-D physics model of the plasma initial phase is developed to guide our 2-D modeling studies.

  7. 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.

  8. 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.

  9. Nonequilibrium diagnostics of plasma thrusters

    SciTech Connect

    Eddy, T.L.; Grandy, J.D.

    1990-01-01

    This paper describes possible techniques by which the state of plasma thruster operation for space propulsion can be determined from a minimum set of experimental data in the laboratory. The kinetic properties of the nonequilibrium plasma plume usually can not be directly related to the observed radiation; hence, appropriate nonequilibrium diagnostic techniques must be employed. A newly developed multithermal, multichemical equilibrium method is discussed that uses measured line emission intensities and N equations to solve for N unknowns. The effect of arbitrarily changing the number of selected N unknowns and how one determines the optimum (minimum) number to be used for a given composition is also presented. The chemical nonequilibrium aspects and the application to molecular species have not yet been published. The important conclusions are that (1) complete thermodynamic systems in nonequilibrium can be described by relatively few variables if appropriate choices and filtering methods are used, (2) a few radiation measurements can yield valid kinetic properties, and (3) the major question in the relations to be used is in the form of the law of mass action. The results are substantiated in the laboratory by additional alternative methods of measurement of some of the kinetic properties. 13 refs., 1 fig.

  10. The solar photon thruster as a terrestrial pole sitter.

    PubMed

    Matloff, Gregory L

    2004-05-01

    Geosynchronous satellites are invisible at high latitudes. A pole-sitting spacecraft would have communication, climate-studies, and near-polar Earth observation applications. We present a pole-sitter based on the solar photon thruster (SPT), a two-sail variant of the solar sail using a large curved collector sail (always normal to the Sun) to direct sunlight against a much smaller thruster. Thrust decreases slower for an SPT than for a conventional sail arrangement as the angle between sunlight and the collector normal increases. An SPT pole-sitter is offset from the terrestrial pole so that a component of Earth gravity balances the solar radiation-pressure component pushing the SPT off station. The component of gravitational attraction of the Earth pulling the spacecraft towards Earth is also balanced by a solar radiation-pressure component. Results are presented for 80-100% collector/thruster reflectivities. For a spacecraft areal mass thickness of 0.002 kg/m(2), collector and thruster reflectivities of 0.9, the SPT can be situated above latitude 45 degrees at a distance of approximately 60 Earth radii. An SPT pole sitter would be affected by lunar perturbation, which can be compensated for by an on-board rocket thruster producing 2 x 10(-6) g acceleration, a second SPT thruster sail thrusting against the influence of the Moon, or by directing a microwave beam against the spacecraft. Since an SPT pole sitter is in a position rather than an orbit, the effect of terrestrial gravitation limits the size and design of the payload package, which limits terrestrial target resolution.

  11. Laboratory-Model Integrated-System FARAD Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, K.A.; Best, S.; Miller, R.; Rose, M.F.; Owens, T.

    2008-01-01

    Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a plasma current sheet in propellant located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current with an induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster [1,2] is a type of pulsed inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate preionization mechanism in this manner allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those found in previous pulsed inductive accelerators like the Pulsed Inductive Thruster (PIT). In a previous paper [3], the authors presented a basic design for a 100 J/pulse FARAD laboratory-version thruster. The design was based upon guidelines and performance scaling parameters presented in Refs. [4, 5]. In this paper, we expand upon the design presented in Ref. [3] by presenting a fully-assembled and operational FARAD laboratory-model thruster and addressing system and subsystem-integration issues (concerning mass injection, preionization, and acceleration) that arose during assembly. Experimental data quantifying the operation of this thruster, including detailed internal plasma measurements, are presented by the authors in a companion paper [6]. The thruster operates by first injecting neutral gas over the face of a flat, inductive acceleration coil and at some later time preionizing the gas. Once the gas is preionized current is passed through the acceleration coil, inducing a plasma current sheet in the propellant that is accelerated away from the coil through electromagnetic interaction with the time-varying magnetic field

  12. Scaling and Systems Considerations in Pulsed Inductive Thrusters

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.

    2007-01-01

    Performance scaling in pulsed inductive thrusters is discussed in the context of previous experimental studies and modeling results. Two processes, propellant ionization and acceleration, are interconnected where overall thruster performance and operation are concerned, but they are separated here to gain physical insight into each process and arrive at quantitative criteria that should be met to address or mitigate inherent inductive thruster difficulties. The effects of preionization in lowering the discharge energy requirements relative to a case where no preionization is employed, and in influencing the location of the initial current sheet, are described. The relevant performance scaling parameters for the acceleration stage are reviewed, emphasizing their physical importance and the numerical values required for efficient acceleration. The scaling parameters are then related to the design of the pulsed power train providing current to the acceleration stage. The impact of various choices in pulsed power train and circuit topology selection are reviewed, paying special attention to how these choices mitigate or exacerbate switching, lifetime, and power consumption issues.

  13. Operational Characteristics of a Low-Energy FARAD Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.; Rose, M. Frank; Miller, Robert

    2008-01-01

    Data from a 100 J per pulse electrodeless accelerator employing pulsed RF-preionization are presented to gain insight into the accelerator's operating characteristics. The data suggest that the propellant distribution is highly unoptimized, with most of the gas inaccessible to the discharge and the remainder mostly concentrated at the inner radius of the coil. The pulsed RF-preionization discharge produces a visible plasma, but like the gas distribution it mostly appears concentrated at the inner radius of the thruster. Magnetic field probes in the discharge point to a current sheet that is not magnetically impermeable. These data also exhibit signs of nonrepeatability, and time-integrated discharge photography shows signs of spatial nonuniformity in both the radial and azimuthal directions. Terminal voltage measurements on the two capacitor banks of the thruster do not exhibit the asymmetric nature (in time) typically associated with an efficient pulsed plasma accelerator. Based on the experimental evidence, the poor performance of the thruster is thought to be due to insufficient preionization, which at these low discharge energy levels severely limits the ability of the main current pulse to couple with and effectively accelerate the propellant.

  14. Measurement and Characterization of the Acceleration Environment on Board the Space Station

    NASA Technical Reports Server (NTRS)

    Baugher, Charles R. (Editor)

    1990-01-01

    This workshop provides a comprehensive overview of the work and status of each of these areas to provide a basis for establishing a systematic approach to the challenge of avoiding these difficulties during the Space Station era of materials experimentation. The discussions were arranged in the order of: the scientific understanding of the requirements for a micro-gravity environment, a history of acceleration measurements on spacecraft, the state of accelerometer technology, and the current understanding of the predicted Space Station environment.

  15. Complex System for Ground-Based and Accelerated Simulation of Six Extremal Space Factors (KIFK)

    NASA Astrophysics Data System (ADS)

    Abraimov, V. V.; Kolybaev, L. K.; Verkhovtseva, E. T.; Nekludov, I. M.; Rybalko, V. F.; Borts, B. V.

    This work was aimed at 1. Development and construction of a complex SF simulator for simultaneous and accelerated simulation of six SF (unparalleled in the GUS and ESA countries): artificial Sun radiation (200..2500 nm) + VUV radiation (5..200 nm) + proton (p+) and electron (e-) radiation (50..200 keV) + vacuum (10-6 Torr) + thermocycling (4.2..400 K). 2.Development of new physical techniques of accelerated laboratory simulation of the basic space factors adequate to natural factors influencing materials, components, units and scale models of spacecraft. In our opinion, this work solves a number of topical problems of ground-based simulation of space factors. 1.Simultaneous impact of six SF on spacecraft materials and models. 2.Simulation of the complete electromagnetic solar radiation (including VUV) spectrum in the wavelength range 5..2500 nm. 3.Accelerated simulation of the electron and proton flows of the Earth's radiation belts (the energy 50..200 keV, intensity 1012 particle/cm2s), i.e. with the acceleration coefficient 500..1000. 4.Accelerated simulation of the VUV component of the Sun's spectrum with the maximum intensity 3000 erg/cm2s) (i.e., the acceleration coefficient 300) 5.Simultaneous thermocycling of test objects in a wide range of temperatures 4.2..400 K in the vacuum of 10-6 Torr.

  16. Magnetically-conformed, Variable Area Discharge Chamber for Hall Thruster, and Method

    NASA Technical Reports Server (NTRS)

    Hofer, Richard R. (Inventor)

    2013-01-01

    The invention is a Hall thruster that incorporates a discharge chamber having a variable area channel including an ionization zone, a transition region, and an acceleration zone. The variable area channel is wider through the acceleration zone than through the ionization zone. An anode is located in a vicinity of the ionization zone and a cathode is located in a vicinity of the acceleration zone. The Hall thruster includes a magnetic circuit which is capable of forming a local magnetic field having a curvature within the transition region of the variable area channel whereby the transition region conforms to the curvature of the local magnetic field. The Hall thruster optimizes the ionization and acceleration efficiencies by the combined effects of the variable area channel and magnetic conformity.

  17. 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

  18. High Voltage Hall Accelerator Propulsion System Development for NASA Science Missions

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hani; Haag, Thomas; Huang, Wensheng; Shastry, Rohit; Pinero, Luis; Peterson, Todd; Dankanich, John; Mathers, Alex

    2013-01-01

    NASA Science Mission Directorates In-Space Propulsion Technology Program is sponsoring the development of a 3.8 kW-class engineering development unit Hall thruster for implementation in NASA science and exploration missions. NASA Glenn Research Center and Aerojet are developing a high fidelity high voltage Hall accelerator (HiVHAc) thruster that can achieve specific impulse magnitudes greater than 2,700 seconds and xenon throughput capability in excess of 300 kilograms. Performance, plume mappings, thermal characterization, and vibration tests of the HiVHAc engineering development unit thruster have been performed. In addition, the HiVHAc project is also pursuing the development of a power processing unit (PPU) and xenon feed system (XFS) for integration with the HiVHAc engineering development unit thruster. Colorado Power Electronics and NASA Glenn Research Center have tested a brassboard PPU for more than 1,500 hours in a vacuum environment, and a new brassboard and engineering model PPU units are under development. VACCO Industries developed a xenon flow control module which has undergone qualification testing and will be integrated with the HiVHAc thruster extended duration tests. Finally, recent mission studies have shown that the HiVHAc propulsion system has sufficient performance for four Discovery- and two New Frontiers-class NASA design reference missions.

  19. High voltage hall accelerator propulsion system development for NASA science missions

    NASA Astrophysics Data System (ADS)

    Kamhawi, Hani; Haag, Thomas; Huang, Wensheng; Shastry, Rohit; Pinero, Luis; Peterson, Todd; Dankanich, John; Mathers, Alex

    NASA Science Mission Directorate's In-Space Propulsion Technology Program is sponsoring the development of a 3.8 kW-class engineering development unit Hall thruster for implementation in NASA science and exploration missions. NASA Glenn Research Center and Aerojet are developing a high fidelity high voltage Hall accelerator (HiVHAc) thruster that can achieve specific impulse magnitudes greater than 2,700 seconds and xenon throughput capability in excess of 300 kilograms. Performance, plume mappings, thermal characterization, and vibration tests of the HiVHAc engineering development unit thruster have been performed. In addition, the HiVHAc project is also pursuing the development of a power processing unit (PPU) and xenon feed system for integration with the HiVHAc engineering development unit thruster. Colorado Power Electronics and NASA Glenn Research Center have tested a brassboard PPU for more than 1,500 hours in a vacuum environment, and a new brassboard and engineering model PPU units are under development. VACCO Industries developed a xenon flow control module which has undergone qualification testing and will be integrated with the HiVHAc thruster extended duration tests. Finally, recent mission studies have shown that the HiVHAc propulsion system has sufficient performance for four Discovery- and two New Frontiers-class NASA design reference missions.

  20. On Spin-Statistics and Bogoliubov Transformations in Flat Space-Time with Acceleration Conditions

    NASA Astrophysics Data System (ADS)

    Good, Michael R. R.

    2013-01-01

    A single real scalar field of spin zero obeying the Klein-Gordon equation in flat space-time under external conditions is considered in the context of the spin-statistics connection. An imposed accelerated boundary on the field is made to become, in the far future, (1) asymptotically inertial and (2) asymptotically noninertial (with an infinite acceleration). The constant acceleration Unruh effect is also considered. The systems involving nontrivial Bogoliubov transformations contain dynamics which point to commutation relations. Particles described by in-modes obey the same statistics as particles described by out-modes. It is found in the nontrivial systems that the spin-statistics connection can be manifest from the acceleration. The equation of motion for the boundary which forever emits thermal radiation is revealed.

  1. Critical role of blockage ratio for flame acceleration in channels with tightly spaced obstacles

    NASA Astrophysics Data System (ADS)

    Ugarte, Orlando J.; Bychkov, Vitaly; Sadek, Jad; Valiev, Damir; Akkerman, V'yacheslav

    2016-09-01

    A conceptually laminar mechanism of extremely fast flame acceleration in obstructed channels, identified by Bychkov et al. ["Physical mechanism of ultrafast flame acceleration," Phys. Rev. Lett. 101, 164501 (2008)], is further studied by means of analytical endeavors and computational simulations of compressible hydrodynamic and combustion equations. Specifically, it is shown how the obstacles length, distance between the obstacles, channel width, and thermal boundary conditions at the walls modify flame propagation through a comb-shaped array of parallel thin obstacles. Adiabatic and isothermal (cold and preheated) side walls are considered, obtaining minor difference between these cases, which opposes the unobstructed channel case, where adiabatic and isothermal walls provide qualitatively different regimes of flame propagation. Variations of the obstructed channel width also provide a minor influence on flame propagation, justifying a scale-invariant nature of this acceleration mechanism. In contrast, the spacing between obstacles has a significant role, although it is weaker than that of the blockage ratio (defined as the fraction of the channel blocked by obstacles), which is the key parameter of the problem. Evolution of the burning velocity and the dependence of the flame acceleration rate on the blockage ratio are quantified. The critical blockage ratio, providing the limitations for the acceleration mechanism in channels with comb-shaped obstacles array, is found analytically and numerically, with good agreement between both approaches. Additionally, this comb-shaped obstacles-driven acceleration is compared to finger flame acceleration and to that produced by wall friction.

  2. Effect of Inductive Coil Geometry on the Operating Characteristics of an Inductive Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Hallock, Ashley K.; Polzin, Kurt A.; Kimberlin, Adam C.; Perdue, Kevin A.

    2012-01-01

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

  3. Transverse space charge effect calculation in the Synergia accelerator modeling toolkit

    SciTech Connect

    Okonechnikov, Konstantin; Amundson, James; Macridin, Alexandru; /Fermilab

    2009-09-01

    This paper describes a transverse space charge effect calculation algorithm, developed in the context of accelerator modeling toolkit Synergia. The introduction to the space charge problem and the Synergia modeling toolkit short description are given. The developed algorithm is explained and the implementation is described in detail. As a result of this work a new space charge solver was developed and integrated into the Synergia toolkit. The solver showed correct results in comparison to existing Synergia solvers and delivered better performance in the regime where it is applicable.

  4. Power Reduction of the Air-Breathing Hall-Effect Thruster

    NASA Astrophysics Data System (ADS)

    Kim, Sungrae

    Electric propulsion system is spotlighted as the next generation space propulsion system due to its benefits; one of them is specific impulse. While there are a lot of types in electric propulsion system, Hall-Effect Thruster, one of electric propulsion system, has higher thrust-to-power ratio and requires fewer power supplies for operation in comparison to other electric propulsion systems, which means it is optimal for long space voyage. The usual propellant for Hall-Effect Thruster is Xenon and it is used to be stored in the tank, which may increase the weight of the thruster. Therefore, one theory that uses the ambient air as a propellant has been proposed and it is introduced as Air-Breathing Hall-Effect Thruster. Referring to the analysis on Air-Breathing Hall-Effect Thruster, the goal of this paper is to reduce the power of the thruster so that it can be applied to real mission such as satellite orbit adjustment. To reduce the power of the thruster, two assumptions are considered. First one is changing the altitude for the operation, while another one is assuming the alpha value that is electron density to ambient air density. With assumptions above, the analysis was done and the results are represented. The power could be decreased to 10s˜1000s with the assumptions. However, some parameters that do not satisfy the expectation, which would be the question for future work, and it will be introduced at the end of the thesis.

  5. Accelerated simulation study of space charge effects in quadrupole ion traps using GPU techniques.

    PubMed

    Xiong, Xingchuang; Xu, Wei; Fang, Xiang; Deng, Yulin; Ouyang, Zheng

    2012-10-01

    Space charge effects play important roles in the performance of various types of mass analyzers. Simulation of space charge effects is often limited by the computation capability. In this study, we evaluate the method of using graphics processing unit (GPU) to accelerate ion trajectory simulation. Simulation using GPU has been compared with multi-core central processing unit (CPU), and an acceleration of about 390 times have been obtained using a single computer for simulation of up to 10(5) ions in quadrupole ion traps. Characteristics of trapped ions can be investigated at detailed levels within a reasonable simulation time. Space charge effects on the trapping capacities of linear and 3D ion traps, ion cloud shapes, ion motion frequency shift, mass spectrum peak coalescence effects between two ion clouds of close m/z are studied with the ion trajectory simulation using GPU.

  6. Dual-throat thruster thermal model

    NASA Technical Reports Server (NTRS)

    Ewen, R. L.; Obrien, C. J.; Matthews, L. W.

    1986-01-01

    The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

  7. Characterization of 8-cm engineering model thruster

    NASA Technical Reports Server (NTRS)

    Williamson, W. S.

    1984-01-01

    Development of 8 cm ion thruster technology which was conducted in support of the Ion Auxiliary Propulsion System (IAPS) flight contract (Contract NAS3-21055) is discussed. The work included characterization of thruster performance, stability, and control; a study of the effects of cathode aging; environmental qualification testing; and cyclic lifetesting of especially critical thruster components.

  8. 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).

  9. The effects of magnetic nozzle configurations on plasma thrusters

    NASA Technical Reports Server (NTRS)

    York, Thomas M.

    1990-01-01

    Plasma thrusters have been operated at power levels from 10 kw to 0.1 MW. When these devices have had magnetic fields applied to them which form a nozzle configuration for the expanding plasma, they have shown marked increases in exhaust velocity which is in direct proportion to the magnitude of the applied field. Further, recent results have shown that electrode erosion may be influenced by applied magnetic fields. This research effort is directed to the experimental and computational study of the effects of applied magnetic field nozzles in the acceleration of plasma flows. Plasma source devices which eliminate the plasma interaction in normal thrusters are studied as most basic. Normal thruster configurations were studied without applied fields and with applied magnetic nozzle fields. Unique computational studies utilize existing codes which accurately include transport processes. Unique diagnostic studies supported the experimental studies to generate new data. Both computation and diagnostics were combined to indicate the physical mechanisms and transport properties that are operative in order to allow scaling and accurate prediction of thruster performance.

  10. The effects of magnetic nozzle configurations on plasma thrusters

    NASA Technical Reports Server (NTRS)

    York, Thomas M.

    1989-01-01

    Plasma thrusters have been operated at power levels from 10kW to 0.1MW. When these devices have had magnetic fields applied to them which form a nozzle configuration for the expanding plasma, they have shown marked increases in exhaust velocity which is in direct proportion to the magnitude of the applied field. Further, recent results have shown that electrode erosion may be influenced by applied magnetic fields. This research is directed to the experimental and computational study of the effects of applied magnetic field nozzles in the acceleration of plasma flows. Plasma source devices which eliminate the plasma interaction in normal thrusters are studied as most basic. Normal thruster configurations will be studied without applied fields and with applied magnetic nozzle fields. Unique computational studies will utilize existing codes which accurately include transport processes. Unique diagnostic studies will support the experimental studies to generate new data. Both computation and diagnostics will be combined to indicate the physical mechanisms and transport properties that are operative in order to allow scaling and accurate prediction of thruster performance.

  11. Power Electronics Development for the SPT-100 Thruster

    NASA Technical Reports Server (NTRS)

    Hamley, John A.; Hill, Gerald M.; Sankovic, John M.

    1994-01-01

    Russian electric propulsion technologies have recently become available on the world market. Of significant interest is the Stationary Plasma Thruster (SPT) which has a significant flight heritage in the former Soviet space program. The SPT has performance levels of up to 1600 seconds of specific impulse at a thrust efficiency of 0.50. Studies have shown that this level of performance is well suited for stationkeeping applications, and the SPT-100, with a 1.35 kW input power level, is presently being evaluated for use on Western commercial satellites. Under a program sponsored by the Innovative Science and Technology Division of the Ballistic Missile Defense Organization, a team of U.S. electric propulsion specialists observed the operation of the SPT-100 in Russia. Under this same program, power electronics were developed to operate the SPT-100 to characterize thruster performance and operation in the U.S. The power electronics consisted of a discharge, cathode heater, and pulse igniter power supplies to operate the thruster with manual flow control. A Russian designed matching network was incorporated in the discharge supply to ensure proper operation with the thruster. The cathode heater power supply and igniter were derived from ongoing development projects. No attempts were made to augment thruster electromagnet current in this effort. The power electronics successfully started and operated the SPT-100 thruster in performance tests at NASA Lewis, with minimal oscillations in the discharge current. The efficiency of the main discharge supply was measured at 0.92, and straightforward modifications were identified which could increase the efficiency to 0.94.

  12. 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.

  13. 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

  14. 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.

  15. 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.

  16. Hybrid-PIC Modeling of a High-Voltage, High-Specific-Impulse Hall Thruster

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

    The primary life-limiting mechanism of Hall thrusters is the sputter erosion of the discharge channel walls by high-energy propellant ions. Because of the difficulty involved in characterizing this erosion experimentally, many past efforts have focused on numerical modeling to predict erosion rates and thruster lifespan, but those analyses were limited to Hall thrusters operating in the 200-400V discharge voltage range. Thrusters operating at higher discharge voltages (V(sub d) >= 500 V) present an erosion environment that may differ greatly from that of the lower-voltage thrusters modeled in the past. In this work, HPHall, a well-established hybrid-PIC code, is used to simulate NASA's High-Voltage Hall Accelerator (HiVHAc) at discharge voltages of 300, 400, and 500V as a first step towards modeling the discharge channel erosion. It is found that the model accurately predicts the thruster performance at all operating conditions to within 6%. The model predicts a normalized plasma potential profile that is consistent between all three operating points, with the acceleration zone appearing in the same approximate location. The expected trend of increasing electron temperature with increasing discharge voltage is observed. An analysis of the discharge current oscillations shows that the model predicts oscillations that are much greater in amplitude than those measured experimentally at all operating points, suggesting that the differences in oscillation amplitude are not strongly associated with discharge voltage.

  17. iSat Surface Charging and Thruster Plume Interactions Analysis

    NASA Technical Reports Server (NTRS)

    Parker, L. Neergaard; Willis, E. M.; Minow, J. I.

    2016-01-01

    Characterizing the electromagnetic interaction of a satellite in low Earth, high inclination orbit with the space plasma environment and identifying viable charging mitigation strategies is a critical mission design task. High inclination orbits expose the vehicle to auroral charging environments that can potentially charge surfaces to kilovolt potentials and electric thruster propulsion systems will interact with the ambient plasma environment throughout the orbit. NASA is designing the Iodine Satellite (iSAT) cubesat mission to demonstrate operations of an iodine electric thruster system. The spacecraft will be deployed as a secondary payload from a launch vehicle which has not yet been identified so the program must plan for the worst case environments over a range of orbital inclinations. We will first present results from a NASA and Air Force Charging Analyzer Program (Nascap) -2k surface charging calculation used to evaluate the effects of auroral charging on the spacecraft and to provide the charging levels at other locations in orbit for a thruster plume interaction analysis for the iSAT mission. We will then discuss results from the thruster interactions analysis using the Electric Propulsion Interactions Code (EPIC) with inputs from Nascap-2k. The results of these analyses are being used by the iSAT program to better understand how their spacecraft will interact with the space plasma environment in the range of environments that could be encountered when the final mission orbit is selected.

  18. Space acceleration measurement system description and operations on the First Spacelab Life Sciences Mission

    NASA Technical Reports Server (NTRS)

    Delombard, Richard; Finley, Brian D.

    1991-01-01

    The Space Acceleration Measurement System (SAMS) project and flight units are briefly described. The SAMS operations during the STS-40 mission are summarized, and a preliminary look at some of the acceleration data from that mission are provided. The background and rationale for the SAMS project is described to better illustrate its goals. The functions and capabilities of each SAMS flight unit are first explained, then the STS-40 mission, the SAMS's function for that mission, and the preparation of the SAMS are described. Observations about the SAMS operations during the first SAMS mission are then discussed. Some sample data are presented illustrating several aspects of the mission's microgravity environment.

  19. A 2000-Hour Durability Test of a 5-Centimeter Diameter Mercury Bombardment Ion Thruster

    NASA Technical Reports Server (NTRS)

    Nakanishi, S.; Finke, R. G.

    1972-01-01

    A 2000-hour durability test of a modified Hughes SIT-5 (Structurally Integrated Thruster, 5 cm) was conducted at the Lewis Research Center. The thruster operated with a translating screen thrust vector grid locked in position for 10 deg beam deflection. The test was essentially continuous except for seven stoppages of beam current. The neutralizer keeper voltage and thruster floating potential increased slightly with time. Performance profiles and maps of thruster characteristics were obtained at 453 and 2023 hours into the test. Overall efficiency was nearly constant at 31 - 32 percent, and operating characteristics were similar at both points in the test. A post-shutdown inspection showed negligible erosion damage to the accelerator and cathode baffle. Some erosion was found in the aperture of the neutralizer cathode.

  20. Three classes of space-time with a uniformly accelerating and nonrotating gravitational source

    NASA Astrophysics Data System (ADS)

    Ishikawa, K.-I.; Miyashita, T.

    1983-11-01

    In his study of Einstein spaces, Petrov (1969) has classified curvature tensors algebraically. The present investigation is concerned with the type-D empty spaces which possess a degenerate principal null vector tangent to expanding and nonrotating null geodesic congruences. These spaces have been represented by the C metric. Ishikawa and Miyashita (1982) have conducted a classification of the type-D empty spaces. In the current study, the physical properties of space-time of the three classes are explored. In a representation of the metric forms of the three classes, use is made of the Eddington-Finkelstein type of null coordinates. It is shown that the considered classes can be interpreted as a gravitational field of a uniformly accelerating source which moves with a velocity which is equal to that of light, or either slower or faster than it.

  1. Progress on the Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

    Martin, Adam; Eskridge, Richard; Fimognari, Peter; Koelfgen, Syri

    2005-01-01

    A plasmoid, also called a compact toroid, is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic field (Bp and Bt, respectively). An object with Bp/Bt much greater than 1 is called a Field Reverse Configuration (FRC); if Bp = Bt, it is called a Spheromak. The plasmoid thruster is a pulsed inductive device which operates by repetitively producing plasmoids that are accelerated and ejected at high velocity. As the process is inductive, this thruster avoids the problem of electrode erosion. Also, the magnetic structure of the plasmoid should suppress thermal and mass losses to the wall, and improve detachment of the plasma exhaust from the thruster. This concept should be capable of producing an Isp of 5,000 seconds and greater, with thrust densities of order 10(exp 5) N/sq m. The plasmoid thruster consists chiefly of a conical theta-pinch coil. Propellant is introduced onto a bias magnetic field, produced by an auxiliary coil, and is then pre-ionized, freezing in the magnetic field. The theta-pinch coil is then energized producing a field aligned anti- parallel to the bias field. The reversed field reconnects with the bias field to form the plasmoid. The magnetic pressure of the reversed field accelerates the plasmoid out of the thruster . A series of experiments have been conducted on the PTX device, which consisted of a single turn conical theta-pinch coil, driven by a 560 nF, 35 kV capacitor bank, which rang at a frequency of 500 kHz, and served all three functions required for formation: pre-ionization, bias field loading, and field reversal. Initial ionization was found to occur in an annular region at the exit plane of the coil, and was found to be reproducible with a variety of gases, including H2, D2, Ar, and an H2/N2 mixture (75% / 25%). A fast gas valve for injecting propellant has been tested, as well as a ringing pre-ionization circuit (operating at 5

  2. Accelerated multiscale space-time finite element simulation and application to high cycle fatigue life prediction

    NASA Astrophysics Data System (ADS)

    Zhang, Rui; Wen, Lihua; Naboulsi, Sam; Eason, Thomas; Vasudevan, Vijay K.; Qian, Dong

    2016-08-01

    A multiscale space-time finite element method based on time-discontinuous Galerkin and enrichment approach is presented in this work with a focus on improving the computational efficiencies for high cycle fatigue simulations. While the robustness of the TDG-based space-time method has been extensively demonstrated, a critical barrier for the extensive application is the large computational cost due to the additional temporal dimension and enrichment that are introduced. The present implementation focuses on two aspects: firstly, a preconditioned iterative solver is developed along with techniques for optimizing the matrix storage and operations. Secondly, parallel algorithms based on multi-core graphics processing unit are established to accelerate the progressive damage model implementation. It is shown that the computing time and memory from the accelerated space-time implementation scale with the number of degree of freedom N through ˜ O(N^{1.6}) and ˜ O(N), respectively. Finally, we demonstrate the accelerated space-time FEM simulation through benchmark problems.

  3. Acceleration of color computer-generated hologram from RGB-D images using color space conversion

    NASA Astrophysics Data System (ADS)

    Hiyama, Daisuke; Shimobaba, Tomoyoshi; Kakue, Takashi; Ito, Tomoyoshi

    2015-04-01

    We report acceleration of color computer-generated holograms (CGHs) from three dimensional (3D) scenes that are expressed as RGB and depth (D) images. These images are captured by a depth camera and depth buffer of 3D graphics library. RGB and depth images preserve color and depth information of 3D scene, respectively. Then we can regard them as two-dimensional (2D) section images along the depth direction. In general, convolution-based diffraction such as the angular spectrum method is used in calculating CGHs from the 2D section images. However, it takes enormous amount of time because of multiple diffraction calculations. In this paper, we first describe 'band-limited double-step Fresnel diffraction (BL-DSF)' which can accelerate the diffraction calculation than convolution-based diffraction. Next, we describe acceleration of color CGH using color space conversion. Color CGHs are generally calculated on RGB color space; however, we need to perform the same calculations for each color component repeatedly, so that computational cost of color CGH calculation is three times as that of monochrome CGH calculation. Instead, we use YCbCr color space because the 2D section images on YCbCr color space can be down-sampled without deterioration of the image quality.

  4. 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.

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

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  6. Hubble Space Telescope Program on STS-95 Supported by Space Acceleration Measurement System for Free Flyers

    NASA Technical Reports Server (NTRS)

    Kacpura, Thomas J.

    2000-01-01

    John Glenn's historic return to space was a primary focus of the STS 95 space shuttle mission; however, the 83 science payloads aboard were the focus of the flight activities. One of the payloads, the Hubble Space Telescope Orbital System Test (HOST), was flown in the cargo bay by the NASA Goddard Space Flight Center. It served as a space flight test of upgrade components for the telescope before they are installed in the shuttle for the next Hubble Space Telescope servicing mission. One of the upgrade components is a cryogenic cooling system for the Near Infrared Camera and Multi-Object Spectrometer (NICMOS). The cooling is required for low noise in the receiver's sensitive electronic instrumentation. Originally, a passive system using dry ice cooled NICMOS, but the ice leaked away and must be replaced. The active cryogenic cooler can provide the cold temperatures required for the NICMOS, but there was a concern that it would create vibrations that would affect the fine pointing accuracy of the Hubble platform.

  7. Stationary Plasma Thruster Ion Velocity Distribution

    NASA Technical Reports Server (NTRS)

    Manzella, David H.

    1994-01-01

    A nonintrusive velocity diagnostic based on laser induced fluorescence of the 5d4F(5/2)-6p4D(5/2) singly ionized xenon transition was used to interrogate the exhaust of a 1.5 kW Stationary Plasma Thruster (SPT). A detailed map of plume velocity vectors was obtained using a simplified, cost-effective, nonintrusive, semiconductor laser based scheme. Circumferential velocities on the order of 250 m/s were measured which implied induced momentum torques of approximately 5 x 10(exp -2) N-cm. Axial and radial velocities were evaluated one mm downstream of the cathode at several locations across the width of the annular acceleration channel. Radial velocities varied linearly with radial distance. A maximum radial velocity of 7500 m/s was measured 8 mm from the center of the channel. Axial velocities as large as 16,500 m/s were measured.

  8. A linear accelerator in the space: The beam experiment aboard rocket

    SciTech Connect

    O'Shea, P.G.; Butler, T.A.; Lynch, M.T.; McKenna, K.F.; Pongratz, M.B.

    1990-01-01

    On July 13, 1989 the BEAM experiment Aboard Rocket (BEAR) linear accelerator was successfully launched and operated in space. The flight demonstrated that a neutral hydrogen beam could be successfully propagated in an exoatmospheric environment. The accelerator, which was the result of an extensive collaboration between Los Alamos National Laboratory and industrial partners, was designed to produce a 10 mA (equivalent), 1 MeV neutral hydrogen beam in 50 {mu}s pulses at 5 Hz. The major components were a 30 keV H{sup {minus}} injector a 1 MeV radio frequency quadrupole, two 425 Mhz RF amplifiers, a gas cell neutralizer, beam optics, vacuum system and controls. The design was strongly constrained by the need for a lightweight rugged system that would survive the rigors of launch and operate autonomously. Following the flight the accelerator was recovered and operated again on the laboratory. 6 figs., 2 tabs.

  9. Development of and flight results from the Space Acceleration Measurement System (SAMS)

    NASA Technical Reports Server (NTRS)

    Delombard, Richard; Finley, Brian D.; Baugher, Charles R.

    1992-01-01

    Described here is the development of and the flight results from the Space Acceleration Measurement System (SAMS) flight units used in the Orbiter middeck, Spacelab module, and the Orbitercargo bay. The SAMS units are general purpose microgravity accelerometers designed to support a variety of science experiments with microgravity acceleration measurements. A total of six flight units have been fabricated; four for use in the Orbiter middeck and Spacelab module, and two for use in the Orbiter cargo bay. The design of the units is briefly described. The initial two flights of SAMS units on STS-40 (June 1991) and STS-43 (August 1991) resulted in 371 megabytes and 2.6 gigabytes of data respectively. Analytical techniques developed to examine this quantity of acceleration data are described and sample plots of analyzed data are illustrated. Future missions for the SAMS units are listed.

  10. Particle-in-cell/accelerator code for space-charge dominated beam simulation

    2012-05-08

    Warp is a multidimensional discrete-particle beam simulation program designed to be applicable where the beam space-charge is non-negligible or dominant. It is being developed in a collaboration among LLNL, LBNL and the University of Maryland. It was originally designed and optimized for heave ion fusion accelerator physics studies, but has received use in a broader range of applications, including for example laser wakefield accelerators, e-cloud studies in high enery accelerators, particle traps and other areas.more » At present it incorporates 3-D, axisymmetric (r,z) planar (x-z) and transverse slice (x,y) descriptions, with both electrostatic and electro-magnetic fields, and a beam envelope model. The code is guilt atop the Python interpreter language.« less

  11. Particle-in-cell/accelerator code for space-charge dominated beam simulation

    SciTech Connect

    2012-05-08

    Warp is a multidimensional discrete-particle beam simulation program designed to be applicable where the beam space-charge is non-negligible or dominant. It is being developed in a collaboration among LLNL, LBNL and the University of Maryland. It was originally designed and optimized for heave ion fusion accelerator physics studies, but has received use in a broader range of applications, including for example laser wakefield accelerators, e-cloud studies in high enery accelerators, particle traps and other areas. At present it incorporates 3-D, axisymmetric (r,z) planar (x-z) and transverse slice (x,y) descriptions, with both electrostatic and electro-magnetic fields, and a beam envelope model. The code is guilt atop the Python interpreter language.

  12. Design of an Integrated-System FARAD Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, K.A.; Rose, R.F.; Miller, R.; Owens, T.

    2007-01-01

    Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a current s heet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current and the induced magne tic field, The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster is a type of pulsed inductive plasma accelerator in which t he plasma is preionized by a mechanism separate from that used to for m the current sheet and accelerate the gas. Employing a separate preionization mechanism allows for the formation of an inductive current s heet at much lower discharge energies and voltages than those used in previous pulsed inductive accelerators like the Pulsed Inductive Thr uster (PIT). In this paper, we present the design of a benchtop FARAD thruster with all the subsystems (mass injection, preionization, and acceleration) integrated into a single unit. Design of the thruster follows the guidelines and similarity performance parameters presented elsewhere. The system is designed to use the ringing, RF-frequency s ignal produced by a discharging Vector Inversion Generator (VIG) to p reionize the gas. The acceleration stage operates on the order of 100 J/pulse and can be driven by several different pulsed powertrains. These include a simple capacitor coupled to the system, a Bernardes and Merryman configuration, and a pulsecompression circuit that takes a temporally broad, low current pulse and transforms it into a short, h igh current pulse. A set of applied magnetic field coils are integrated into the system to guide the preionized propellant as it spreads ov er the face of the inductive acceleration coil. The coils are operate d in a pulsed mode, and the thruster can be operated without using the coils to determine if there is a performance

  13. Measuring test mass acceleration noise in space-based gravitational wave astronomy

    NASA Astrophysics Data System (ADS)

    Congedo, Giuseppe

    2015-03-01

    The basic constituent of interferometric gravitational wave detectors—the test-mass-to-test-mass interferometric link—behaves as a differential dynamometer measuring effective differential forces, comprising an integrated measure of gravity curvature, inertial effects, as well as nongravitational spurious forces. This last contribution is going to be characterized by the LISA Pathfinder mission, a technology precursor of future space-borne detectors like eLISA. Changing the perspective from displacement to acceleration can benefit the data analysis of LISA Pathfinder and future detectors. The response in differential acceleration to gravitational waves is derived for a space-based detector's interferometric link. The acceleration formalism can also be integrated into time delay interferometry by building up the unequal-arm Michelson differential acceleration combination. The differential acceleration is nominally insensitive to the system's free evolution dominating the slow displacement dynamics of low-frequency detectors. Working with acceleration also provides an effective way to subtract measured signals acting as systematics, including the actuation forces. Because of the strong similarity with the equations of motion, the optimal subtraction of systematic signals, known within some amplitude and time shift, with the focus on measuring the noise provides an effective way to solve the problem and marginalize over nuisance parameters. The F statistic, in widespread use throughout the gravitation waves community, is included in the method and suitably generalized to marginalize over linear parameters and noise at the same time. The method is applied to LPF simulator data and, thanks to its generality, can also be applied to the data reduction and analysis of future gravitational wave detectors.

  14. Design of a cusped field thruster for drag-free flight

    NASA Astrophysics Data System (ADS)

    Liu, H.; Chen, P. B.; Sun, Q. Q.; Hu, P.; Meng, Y. C.; Mao, W.; Yu, D. R.

    2016-09-01

    Drag-free flight has played a more and more important role in many space missions. The thrust control system is the key unit to achieve drag-free flight by providing a precise compensation for the disturbing force except gravity. The cusped field thruster has shown a significant potential to be capable of the function due to its long life, high efficiency, and simplicity. This paper demonstrates a cusped field thruster's feasibility in drag-free flight based on its instinctive characteristics and describes a detailed design of a cusped field thruster made by Harbin Institute of Technology (HIT). Furthermore, the performance test is conducted, which shows that the cusped field thruster can achieve a continuously variable thrust from 1 to 20 mN with a low noise and high resolution below 650 W, and the specific impulse can achieve 1800 s under a thrust of 18 mN and discharge voltage of 1000 V. The thruster's overall performance indicates that the cusped field thruster is quite capable of achieving drag-free flight. With the further optimization, the cusped field thruster will exhibit a more extensive application value.

  15. 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.

  16. Simulation of Cascaded Longitudinal-Space-Charge Amplifier at the Fermilab Accelerator Science & Technology (Fast) Facility

    SciTech Connect

    Halavanau, A.; Piot, P.

    2015-12-01

    Cascaded Longitudinal Space Charge Amplifiers (LSCA) have been proposed as a mechanism to generate density modulation over a board spectral range. The scheme has been recently demonstrated in the optical regime and has confirmed the production of broadband optical radiation. In this paper we investigate, via numerical simulations, the performance of a cascaded LSCA beamline at the Fermilab Accelerator Science & Technology (FAST) facility to produce broadband ultraviolet radiation. Our studies are carried out using elegant with included tree-based grid-less space charge algorithm.

  17. Phase-space dynamics of ionization injection in plasma-based accelerators.

    PubMed

    Xu, X L; Hua, J F; Li, F; Zhang, C J; Yan, L X; Du, Y C; Huang, W H; Chen, H B; Tang, C X; Lu, W; Yu, P; An, W; Joshi, C; Mori, W B

    2014-01-24

    The evolution of beam phase space in ionization injection into plasma wakefields is studied using theory and particle-in-cell simulations. The injection process involves both longitudinal and transverse phase mixing, leading initially to a rapid emittance growth followed by oscillation, decay, and a slow growth to saturation. An analytic theory for this evolution is presented and verified through particle-in-cell simulations. This theory includes the effects of injection distance (time), acceleration distance, wakefield structure, and nonlinear space charge forces, and it also shows how ultralow emittance beams can be produced using ionization injection methods.

  18. Video-Bubbles Act as an Accelerator on the International Space Station (ISS)

    NASA Technical Reports Server (NTRS)

    2005-01-01

    Saturday Morning Science, the science of opportunity series of applied experiments and demonstrations, performed aboard the International Space Station (ISS) by Expedition 6 astronaut Dr. Don Pettit, revealed some remarkable findings. Bubbles in a flask of water act as an accelerometer in space. In this video, Pettit demonstrates bubbles moving in a direction opposite to the residual g-vector. The demonstration shows scientists that they may need to consider the direction of the residual accelerations as influenced by the orientation of the ISS for future experiments.

  19. Effects of ionization distribution on plasma beam focusing characteristics in Hall thrusters

    SciTech Connect

    Ning Zhongxi; Liu Hui; Yu Daren; Zhou Zhongxiang

    2011-11-28

    The relationship between ionization distribution and divergence of plasma beam in a Hall thruster is investigated using spectrum and probe methods. Experimental results indicate that the shift of ionization region towards the exit of channel causes the reduction of accelerating field and the enhancement of electron thermal pressure effect, which result in further deviation of equipotential lines to magnetic field lines and further increase in divergence of plasma beam. It is, therefore, suggested that to put the ionization region deep inside the channel and separate it from the acceleration region at the design, and development stage is helpful to improve the plasma beam focusing characteristics of a Hall thruster.

  20. 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.

  1. 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.

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

    NASA Technical Reports Server (NTRS)

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

    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 sputting 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.

  3. Improvement of the low frequency oscillation model for Hall thrusters

    NASA Astrophysics Data System (ADS)

    Wang, Chunsheng; Wang, Huashan

    2016-08-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

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

  5. An Experimental Study of a Pulsed Electromagnetic Plasma Accelerator

    NASA Technical Reports Server (NTRS)

    Thio, Y. C. Francis; Eskridge, Richard; Lee, Mike; Smith, James; Martin, Adam; Markusic, Tom E.; Cassibry, Jason T.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    Experiments are being performed on the NASA Marshall Space Flight Center (MSFC) pulsed electromagnetic plasma accelerator (PEPA-0). Data produced from the experiments provide an opportunity to further understand the plasma dynamics in these thrusters via detailed computational modeling. The detailed and accurate understanding of the plasma dynamics in these devices holds the key towards extending their capabilities in a number of applications, including their applications as high power (greater than 1 MW) thrusters, and their use for producing high-velocity, uniform plasma jets for experimental purposes. For this study, the 2-D MHD modeling code, MACH2, is used to provide detailed interpretation of the experimental data. At the same time, a 0-D physics model of the plasma initial phase is developed to guide our 2-D modeling studies.

  6. Transport calculations and accelerator experiments needed for radiation risk assessment in space.

    PubMed

    Sihver, Lembit

    2008-01-01

    The major uncertainties on space radiation risk estimates in humans are associated to the poor knowledge of the biological effects of low and high LET radiation, with a smaller contribution coming from the characterization of space radiation field and its primary interactions with the shielding and the human body. However, to decrease the uncertainties on the biological effects and increase the accuracy of the risk coefficients for charged particles radiation, the initial charged-particle spectra from the Galactic Cosmic Rays (GCRs) and the Solar Particle Events (SPEs), and the radiation transport through the shielding material of the space vehicle and the human body, must be better estimated Since it is practically impossible to measure all primary and secondary particles from all possible position-projectile-target-energy combinations needed for a correct risk assessment in space, accurate particle and heavy ion transport codes must be used. These codes are also needed when estimating the risk for radiation induced failures in advanced microelectronics, such as single-event effects, etc., and the efficiency of different shielding materials. It is therefore important that the models and transport codes will be carefully benchmarked and validated to make sure they fulfill preset accuracy criteria, e.g. to be able to predict particle fluence, dose and energy distributions within a certain accuracy. When validating the accuracy of the transport codes, both space and ground based accelerator experiments are needed The efficiency of passive shielding and protection of electronic devices should also be tested in accelerator experiments and compared to simulations using different transport codes. In this paper different multipurpose particle and heavy ion transport codes will be presented, different concepts of shielding and protection discussed, as well as future accelerator experiments needed for testing and validating codes and shielding materials. PMID:19205295

  7. Electrostatic ion thruster optics calculations

    NASA Technical Reports Server (NTRS)

    Whealton, John H.; Kirkman, David A.; Raridon, R. J.

    1992-01-01

    Calculations have been performed which encompass both a self-consistent ion source extraction plasma sheath and the primary ion optics including sheath and electrode-induced aberrations. Particular attention is given to the effects of beam space charge, accelerator geometry, and properties of the downstream plasma sheath on the position of the electrostatic potential saddle point near the extractor electrode. The electron blocking potential blocking is described as a function of electrode thickness and secondary plasma processes.

  8. 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

  9. 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.

  10. Advanced-technology 30-cm-diameter mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Beattie, J. R.; Kami, S.

    1982-01-01

    An advanced-technology mercury ion thruster designed for operation at high thrust and high thrust-to-power ratio is described. The laboratory-model thruster employs a highly efficient discharge-chamber design that uses high-field-strength samarium-cobalt magnets arranged in a ring-cusp configuration. Ion extraction is achieved using an advanced three-grid ion-optics assembly which utilizes flexible mounts for supporting the screen, accel, and decel electrodes. Performance results are presented for operation at beam currents in the range from 1 to 5 A. The baseline specific discharge power is shown to be about 125 eV/ion, and the acceptable range of net-to-total accelerating-voltage ratio is shown to be in the range of 0.2-0.8 for beam currents in the range of 1-5 A.

  11. An approach to the parametric design of ion thrusters

    NASA Technical Reports Server (NTRS)

    Wilbur, Paul J.; Beattie, John R.; Hyman, Jay, Jr.

    1988-01-01

    A methodology that can be used to determine which of several physical constraints can limit ion thruster power and thrust, under various design and operating conditions, is presented. The methodology is exercised to demonstrate typical limitations imposed by grid system span-to-gap ratio, intragrid electric field, discharge chamber power per unit beam area, screen grid lifetime, and accelerator grid lifetime constraints. Limitations on power and thrust for a thruster defined by typical discharge chamber and grid system parameters when it is operated at maximum thrust-to-power are discussed. It is pointed out that other operational objectives such as optimization of payload fraction or mission duration can be substituted for the thrust-to-power objective and that the methodology can be used as a tool for mission analysis.

  12. Thrust Stand Measurements of the Conical Theta Pinch FARAD Thruster

    NASA Technical Reports Server (NTRS)

    Hallock, Ashley K.; Polzin, Kurt A.

    2010-01-01

    It is found that the impulse of a pulsed inductive plasma thruster utilizing preionization is maximized for a particular ratio of the stored energy in the capacitor to the injected propellant mass. The fact that the impulse depends on the ratio of the initial stored energy to injected propellant mass agrees with previous current sheet studies, supporting the idea that a Townsend-like breakdown process strongly influences current sheet formation, and in turn, current sheet formation strongly affects the operational efficiency of the device. The optimum in half cone angle of the inductive coil can be explained in terms of a balance between the direct axial acceleration and the radial pinching contribution to thrust. From the trends in these data we conclude that operation at the correct ratio of capacitor energy to propellant mass is essential for efficient operation of pulsed inductive plasma thrusters employing a preionized propellant.

  13. Development of a high-speed, reciprocating electrostatic probe system for Hall thruster interrogation

    NASA Astrophysics Data System (ADS)

    Haas, James M.; Gallimore, Alec D.; McFall, Keith; Spanjers, Greg

    2000-11-01

    The use of electrostatic probes to measure local plasma parameters inside the discharge chamber of a Hall thruster presents significant difficulties. The high-temperature, dense plasma, and Hall current in the accelerating channel heat the probe rapidly causing ablation of probe material, which perturbs thruster operation and reduces probe lifetime. Results are presented which show the extent of perturbation to discharge current, cathode potential, and thrust for the case where probe material is ablated. A simple thermal model of probe material heating is developed and ablation times for a typical probe configuration are presented. Using the results of the thermal model, a high-speed axial reciprocating probe (HARP) system was developed to enable probe survival and reduce thruster perturbations during interrogation of the discharge chamber of a Hall thruster. Results using the HARP system are presented showing a significant reduction in thruster perturbation. The results also indicate that a mechanism other than material ablation is contributing to perturbation of the thruster. Based on emissive probe data, the tungsten conductor appears to provide a low impedance path between magnetic field lines, enhancing electron transport to the anode.

  14. Design of a High-Energy, Two-Stage Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Markusic, T. E.; Thio, Y. C. F.; Cassibry, J. T.; Rodgers, Stephen L. (Technical Monitor)

    2002-01-01

    Design details of a proposed high-energy (approx. 50 kJ/pulse), two-stage pulsed plasma thruster are presented. The long-term goal of this project is to develop a high-power (approx. 500 kW), high specific impulse (approx. 7500 s), highly efficient (approx. 50%),and mechanically simple thruster for use as primary propulsion in a high-power nuclear electric propulsion system. The proposed thruster (PRC-PPT1) utilizes a valveless, liquid lithium-fed thermal plasma injector (first stage) followed by a high-energy pulsed electromagnetic accelerator (second stage). A numerical circuit model coupled with one-dimensional current sheet dynamics, as well as a numerical MHD simulation, are used to qualitatively predict the thermal plasma injection and current sheet dynamics, as well as to estimate the projected performance of the thruster. A set of further modelling efforts, and the experimental testing of a prototype thruster, is suggested to determine the feasibility of demonstrating a full scale high-power thruster.

  15. Quick look report of acceleration measurements on Mir Space Station during Mir-16

    NASA Astrophysics Data System (ADS)

    Delombard, Richard; Rogers, Melissa J. B.

    1995-01-01

    The NASA Microgravity Science and Applications Division (MSAD) sponsors the Space Acceleration Measurement System (SAMS) to support microgravity science experiments with microgravity acceleration measurements. In the past, SAMS was flown exclusively on the NASA Orbiters. MSAD is currently sponsoring science experiments participating in the Shuttle-Mir Science Program in cooperation with the Russians on the Mir space station. Included in the complement of MSAD experiments and equipment is a SAMS unit installed on the Mir space station. On 25 August 1994, the SAMS unit was launched on a Russia Progress vehicle to the Mir space station. The SAMS unit will support science experiments from the U.S. and Russia in a manner similar to the Orbiter missions by measuring the microgravity environments during the experiment operations. In October 1994, the SAMS unit recorded data on Mir for over fifty-three hours in seven different time periods to survey possible locations for future experiments. This report presents a quick look at the data and the microgravity environment during the time periods in which SAMS data were acquired. The Mir acceleration data were examined to ascertain gross attributes of the data. This report does not present an exhaustive examination of all possible activities due to the short time available to prepare this quick look report and due to the absence of complete timeline information during the SAMS recording time periods. Appendices A, B and C provide plots of the SAMS data for an overview of the microgravity environment at the times that data were recorded. Appendix D describes the procedures to access SAMS data by file transfer protocol (ftp) utilizing the internet. Appendix E contains a user comment sheet.

  16. The BMDO Thruster-on-a-Pallet Program

    NASA Technical Reports Server (NTRS)

    Caveny, Leonard H.; Curran, Francis M.; Sankovic, John M.; Allen, Douglas M.; Brophy, John R.; Garner, Charles

    1995-01-01

    The Ballistic Missile Defense Organization sponsors an aggressive program to develop and demonstrate electric propulsion and space power technologies for future missions. This program supports a focused effort to design, fabricate, and space qualify a Russian Hall thruster system-on-a-pallet ready to take advantage of a near-term flight opportunity. The Russian Hall Effect Thruster Technology (RHETT) program will demonstrate an integrated pallet design in late FY95. The program also includes a parallel effort to develop advanced Solar Concentrator Arrays with Refractive Linear Element Technology (SCARLET). This synergistic technology will be demonstrated in a flight experiment this summer on the Comet satellite. This paper provides an overview of the RHETT and SCARLET programs with an emphasis on electric propulsion, recent progress, and near-term program plans.

  17. Flame acceleration and DDT in channels with obstacles: Effect of obstacle spacing

    SciTech Connect

    Gamezo, Vadim N.; Oran, Elaine S.; Ogawa, Takanobu

    2008-10-15

    We study flame acceleration and deflagration-to-detonation transition (DDT) in obstructed channels using 2D reactive Navier-Stokes numerical simulations. The energy release rate for the stoichiometric hydrogen-air mixture is modeled by one-step Arrhenius kinetics. Computations performed for channels with symmetrical and staggered obstacle configurations show two main effects of obstacle spacing S. First, more obstacles per unit length create more perturbations that increase the flame surface area more quickly, and therefore the flame speed grows faster. Second, DDT occurs more easily when the obstacle spacing is large enough for Mach stems to form between obstacles. These two effects are responsible for three different regimes of flame acceleration and DDT observed in simulations: (1) Detonation is ignited when a Mach stem formed by the diffracting shock reflecting from the side wall collides with an obstacle, (2) Mach stems do not form, and the detonation is not ignited, and (3) Mach stems do not form, but the leading shock becomes strong enough to ignite a detonation by direct collision with the top of an obstacle. Regime 3 is observed for small S and involves multiple isolated detonations that appear between obstacles and play a key role in final stages of flame and shock acceleration. For Regime 1 and staggered obstacle configurations, we observe resonance phenomena that significantly reduce the DDT time when S/2 is comparable to the channel width. Effects of imposed symmetry and stochasticity on DDT phenomena are also considered. (author)

  18. Dynamic response of a poroelastic half-space to accelerating or decelerating trains

    NASA Astrophysics Data System (ADS)

    Cao, Zhigang; Boström, Anders

    2013-05-01

    The dynamic response of a fully saturated poroelastic half-space due to accelerating or decelerating trains is investigated by a semi-analytical method. The ground is modeled as a saturated poroelastic half-space and Biot's theory is applied to characterize the soil medium, taking the coupling effects between the soil skeleton and the pore fluid into account. A detailed track system is considered incorporating rails, sleepers and embankment, which are modeled as Euler-Bernoulli beams, an anisotropic Kirchhoff plate, and an elastic layer, respectively. The acceleration or deceleration of the train is simulated by properly choosing the time history of the train speed using Fourier transforms combined with Fresnel integrals in the transformed domain. The time domain results are obtained by the fast Fourier transform (FFT). It is found that the deceleration of moving trains can cause a significant increase to the ground vibrations as well as the excess pore water pressure responses at the train speed 200 km/h. Furthermore, the single-phase elastic soil model would underestimate the vertical displacement responses caused by both the accelerating and decelerating trains at the speed 200 km/h.

  19. Controlled Electron Injection into Plasma Accelerators and SpaceCharge Estimates

    SciTech Connect

    Fubiani, Gwenael G.J.

    2005-09-01

    Plasma based accelerators are capable of producing electron sources which are ultra-compact (a few microns) and high energies (up to hundreds of MeVs) in much shorter distances than conventional accelerators. This is due to the large longitudinal electric field that can be excited without the limitation of breakdown as in RF structures.The characteristic scale length of the accelerating field is the plasma wavelength and for typical densities ranging from 1018 - 1019 cm-3, the accelerating fields and scale length can hence be on the order of 10-100GV/m and 10-40 μm, respectively. The production of quasimonoenergetic beams was recently obtained in a regime relying on self-trapping of background plasma electrons, using a single laser pulse for wakefield generation. In this dissertation, we study the controlled injection via the beating of two lasers (the pump laser pulse creating the plasma wave and a second beam being propagated in opposite direction) which induce a localized injection of background plasma electrons. The aim of this dissertation is to describe in detail the physics of optical injection using two lasers, the characteristics of the electron beams produced (the micrometer scale plasma wavelength can result in femtosecond and even attosecond bunches) as well as a concise estimate of the effects of space charge on the dynamics of an ultra-dense electron bunch with a large energy spread.

  20. 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.

  1. An Acceleration Measurement Capability on International Space Station Supporting Microgravity Science Payloads

    NASA Technical Reports Server (NTRS)

    Sutliff, Thomas J.

    1997-01-01

    The conditions present in low-earth-orbit have enabled various scientific investigations to be conducted. A fundamental reason that space is a useful environment for science is the presence of a decreased effect of gravity while in orbital free-fall. Science investigations are conducted to study phenomena occurring as a result of reductions in buoyancy forces and phenomena enabled as a result of the apparent lack of gravity. The science community requires knowledge of the environment in which their experiments are being conducted. A general purpose measurement system was developed in the 1980's to supply scientific investigators with data of the residual acceleration environment. The Space Acceleration Measurement System, or SAMS, has flown aboard the NASA shuttle fleet over fourteen times, successfully acquiring acceleration measurement data in support of microgravity science experiments. As NASA continues its development of the International Space Station and its associated research payloads, it once again becomes necessary to address the requirement to measure the residual acceleration in support of the planned scientific investigations. As a follow-on to the SAMS project, SAMS-2 was initiated to investigate the requirements and develop the measurement system to support the microgravity science payloads aboard the International Space Station. The characteristics of the payload complement to be supported by SAMS-2 and the different capabilities provided by the station requires a fresh look at the design implementation of a general purpose measurement system. Not only is there a requirement for measurements in support of numerous, potentially simultaneous experiments, but this support is required to be provided for at least ten years of on-orbit operation as compared to a ten to fourteen day mission on shuttle. These differences in fundamental requirements result in the design configuration of the space station-based SAMS-2 being significantly different than that

  2. 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.

  3. 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 in the USA, Europe, Japan, China, and USSR are reviewed with reference to the major flight qualified electric propulsion systems. These include resistojets, ion thrusters, ablative pulsed plasma thrusters, stationary plasma thrusters, pulsed magnetoplasmic thrusters, and arcjets. Evolutionary mission applications are presented for high specific impulse electric thruster systems. The current status of arcjet, ion, and magnetoplasmadynamic thrusters and their associated power processor technologies are summarized.

  4. High-Energy Two-Stage Pulsed Plasma Thruster

    NASA Technical Reports Server (NTRS)

    Markusic, Tom

    2003-01-01

    A high-energy (28 kJ per pulse) two-stage pulsed plasma thruster (MSFC PPT-1) has been constructed and tested. The motivation of this project is to develop a high power (approximately 500 kW), high specific impulse (approximately 10000 s), highly efficient (greater than 50%) thruster for use as primary propulsion in a high power nuclear electric propulsion system. PPT-1 was designed to overcome four negative characteristics which have detracted from the utility of pulsed plasma thrusters: poor electrical efficiency, poor propellant utilization efficiency, electrode erosion, and reliability issues associated with the use of high speed gas valves and high current switches. Traditional PPTs have been plagued with poor efficiency because they have not been operated in a plasma regime that fully exploits the potential benefits of pulsed plasma acceleration by electromagnetic forces. PPTs have generally been used to accelerate low-density plasmas with long current pulses. Operation of thrusters in this plasma regime allows for the development of certain undesirable particle-kinetic effects, such as Hall effect-induced current sheet canting. PPT-1 was designed to propel a highly collisional, dense plasma that has more fluid-like properties and, hence, is more effectively pushed by a magnetic field. The high-density plasma loading into the second stage of the accelerator is achieved through the use of a dense plasma injector (first stage). The injector produces a thermal plasma, derived from a molten lithium propellant feed system, which is subsequently accelerated by the second stage using mega-amp level currents, which eject the plasma at a speed on the order of 100 kilometers per second. Traditional PPTs also suffer from dynamic efficiency losses associated with snowplow loading of distributed neutral propellant. The twostage scheme used in PPT-I allows the propellant to be loaded in a manner which more closely approximates the optimal slug loading. Lithium propellant

  5. Characteristics of 30-centimeter mercury ion thrusters

    NASA Technical Reports Server (NTRS)

    Maloy, J. E.; Poeschel, R. L.; Dugeroff, C. R.

    1981-01-01

    The technology development of the 30 centimeter J series mercury ion thruster for prime propulsion application in solar electric propulsion systems is described. Thruster design is reviewed. A standardized set of test and data recording procedures formulated to allow for the characterization of the J series thruster is described. Characteristics measured are the magnetic baffle characterization, the neutralizer characterization, perveance, the minimum eV/ion measurement, and the electrical and propellant utilization efficiency measurements. Test results are presented.

  6. SAMS Acceleration Measurements on MIR

    NASA Technical Reports Server (NTRS)

    Moskowitz, Milton E.; Hrovat, Kenneth; Finkelstein, Robert; Reckart, Timothy

    1997-01-01

    During NASA Increment 3 (September 1996 to January 1997), about 5 gigabytes of acceleration data were collected by the Space Acceleration Measurement System (SAMS) onboard the Russian Space Station, Mir. The data were recorded on 11 optical disks and were returned to Earth on STS-81. During this time, SAMS data were collected in the Priroda module to support the following experiments: the Mir Structural Dynamics Experiment (MiSDE) and Binary Colloidal Alloy Tests (BCAT). This report points out some of the salient features of the microgravity environment to which these experiments were exposed. Also documented are mission events of interest such as the docked phase of STS-81 operations, a Progress engine burn, attitude control thruster operation, and crew exercise. Also included are a description of the Mir module orientations, and the panel notations within the modules. This report presents an overview of the SAMS acceleration measurements recorded by 10 Hz and 100 Hz sensor heads. Variations in the acceleration environment caused by unique activities such as crew exercise and life-support fans are presented. The analyses included herein complement those presented in previous mission summary reports published by the Principal Investigator Microgravity Services (PIMS) group.

  7. Perception of tilt (somatogravic illusion) in response to sustained linear acceleration during space flight

    NASA Technical Reports Server (NTRS)

    Clement, G.; Moore, S. T.; Raphan, T.; Cohen, B.

    2001-01-01

    During the 1998 Neurolab mission (STS-90), four astronauts were exposed to interaural and head vertical (dorsoventral) linear accelerations of 0.5 g and 1 g during constant velocity rotation on a centrifuge, both on Earth and during orbital space flight. Subjects were oriented either left-ear-out or right-ear-out (Gy centrifugation), or lay supine along the centrifuge arm with their head off-axis (Gz centrifugation). Pre-flight centrifugation, producing linear accelerations of 0.5 g and 1 g along the Gy (interaural) axis, induced illusions of roll-tilt of 20 degrees and 34 degrees for gravito-inertial acceleration (GIA) vector tilts of 27 degrees and 45 degrees , respectively. Pre-flight 0.5 g and 1 g Gz (head dorsoventral) centrifugation generated perceptions of backward pitch of 5 degrees and 15 degrees , respectively. In the absence of gravity during space flight, the same centrifugation generated a GIA that was equivalent to the centripetal acceleration and aligned with the Gy or Gz axes. Perception of tilt was underestimated relative to this new GIA orientation during early in-flight Gy centrifugation, but was close to the GIA after 16 days in orbit, when subjects reported that they felt as if they were 'lying on side'. During the course of the mission, inflight roll-tilt perception during Gy centrifugation increased from 45 degrees to 83 degrees at 1 g and from 42 degrees to 48 degrees at 0.5 g. Subjects felt 'upside-down' during in-flight Gz centrifugation from the first in-flight test session, which reflected the new GIA orientation along the head dorsoventral axis. The different levels of in-flight tilt perception during 0.5 g and 1 g Gy centrifugation suggests that other non-vestibular inputs, including an internal estimate of the body vertical and somatic sensation, were utilized in generating tilt perception. Interpretation of data by a weighted sum of body vertical and somatic vectors, with an estimate of the GIA from the otoliths, suggests that

  8. 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. __________________________________________________

  9. 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.

  10. Vibration isolation technology: Sensitivity of selected classes of space experiments to residual accelerations

    NASA Technical Reports Server (NTRS)

    Alexander, J. Iwan D.; Zhang, Y. Q.; Adebiyi, Adebimpe

    1989-01-01

    Progress performed on each task is described. Order of magnitude analyses related to liquid zone sensitivity and thermo-capillary flow sensitivity are covered. Progress with numerical models of the sensitivity of isothermal liquid zones is described. Progress towards a numerical model of coupled buoyancy-driven and thermo-capillary convection experiments is also described. Interaction with NASA personnel is covered. Results to date are summarized and they are discussed in terms of the predicted space station acceleration environment. Work planned for the second year is also discussed.

  11. PARALLEL 3-D SPACE CHARGE CALCULATIONS IN THE UNIFIED ACCELERATOR LIBRARY.

    SciTech Connect

    D'IMPERIO, N.L.; LUCCIO, A.U.; MALITSKY, N.

    2006-06-26

    The paper presents the integration of the SIMBAD space charge module in the UAL framework. SIMBAD is a Particle-in-Cell (PIC) code. Its 3-D Parallel approach features an optimized load balancing scheme based on a genetic algorithm. The UAL framework enhances the SIMBAD standalone version with the interactive ROOT-based analysis environment and an open catalog of accelerator algorithms. The composite package addresses complex high intensity beam dynamics and has been developed as part of the FAIR SIS 100 project.

  12. 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.

  13. Predicting Hall Thruster Operational Lifetime

    NASA Technical Reports Server (NTRS)

    Manzella, David; Yim, John; Boyd, Iain

    2004-01-01

    A simple analytic model predicted Hall thruster channel erosion based on thruster geometry, operating conditions, and magnetic field configuration. This model relied on a one-dimensional representation of the plasma with a fixed ionization fraction and variable ion energies based on the magnetic field distribution. Sputtering was modeled as the result of elastic scattering of ions by neutrals within the channel. Not all scattered ions and neutrals were assumed to reach the channel walls as a result of additional subsequent scattering events. Incorporating this phenomenon resulted in a greater predicted decrease in erosion rate with time than predicted based only on geometric effects. Results from this model were compared to SPT 100 experimental erosion data.

  14. Design of a Low-Energy FARAD Thruster

    NASA Technical Reports Server (NTRS)

    Polzin, K. A.; Rose, M. F.; Miller, R.; Best, S.; Owens, T.; Dankanich, J.

    2007-01-01

    The design of an electrodeless thruster that relies on a pulsed, rf-assisted discharge and electromagnetic acceleration using an inductive coil is presented. The thruster design is optimized using known performance,scaling parameters, and experimentally-determined design rules, with design targets for discharge energy, plasma exhaust velocity; and thrust efficiency of 100 J/pulse, 25 km/s, and 50%, respectively. Propellant is injected using a high-speed gas valve and preionized by a pulsed-RF signal supplied by a vector inversion generator, allowing for current sheet formation at lower discharge voltages and energies relative to pulsed inductive accelerators that do not employ preionization. The acceleration coil is designed to possess an inductance of at least 700 nH while the target stray (non-coil) inductance in the circuit is 70 nH. A Bernardes and Merryman pulsed power train or a pulse compression power train provide current to the acceleration coil and solid-state components are used to switch both powertrains.

  15. Arcjet thruster research and technology

    NASA Technical Reports Server (NTRS)

    Makel, Darby B.; Cann, Gordon L.

    1988-01-01

    The design, analysis, and performance testing of an advanced lower power arcjet is described. A high impedance, vortex stabilized 1-kw class arcjet has been studied. A baseline research thruster has been built and endurance and performance tested. This advanced arcjet has demonstrated long lifetime characteristics, but lower than expected performance. Analysis of the specific design has identified modifications which should improve performance and maintain the long life time shown by the arcjet.

  16. Plasma thrusters development in France

    NASA Astrophysics Data System (ADS)

    Rolfo, André; Cadiou, Anne; Secheresse, O.; Dumazert, P.; Gounot, V.; Ragot, X.; Mattei, N.; Grassin, T.; Garnero, P.

    2002-07-01

    This paper presents an overview of the FRENCH plasma propulsion activities. The main existing and future projects are described. The field of application of plasma propulsion is the station keeping and the orbit raising of geostationary telecommunication satellites (STENTOR) and the transfer of interplanetary vehicles such as Mars Sample Return. The works done in the frame of the preparation of the first commercial spacecraft as well as the preparation of the future and the associated Research and Technology program are described. The scientific activity supporting the development of Hall thrusters is on-going in the frame of the GDR (Groupement de Recherche) CNRS/CNES/Snecma Moteurs /ONERA on Plasma Propulsion. Several Russian entities are also involved: the MIREA (Moscow Institute of Radioelectronics and Automatics), of the RIAME MAI (Research Institute of Applied Mechanics and Electrodynamics - Moscow Aviation Institute) and of the SPT « father å Professor MOROZOV The industrial development activities are jointly conducted by Snecma Moteurs and Russian manufacturer FAKEL. The future developments are mainly dedicated to the use of electric propulsion for the orbit raising of telecommunications satellites which leads to the development of thrusters with higher thrust than those existing today. Works are also performed to develop and improve the tools necessary to evaluate the plume effects of plasma thrusters.

  17. Stationary Plasma Thruster Plume Emissions

    NASA Technical Reports Server (NTRS)

    Manzella, David H.

    1994-01-01

    The emission spectrum from a xenon plasma produced by a Stationary Plasma Thruster provided by the Ballistic Missile Defense Organization (BMDO) was measured. Approximately 270 individual Xe I, Xe II, and XE III transitions were identified. A total of 250 mW of radiated optical emission was estimated from measurements taken at the thruster exit plane. There was no evidence of erosion products in the emission signature. Ingestion and ionization of background gas at elevated background pressure was detected. The distribution of excited states could be described by temperatures ranging from fractions of 1 eV to 4 eV with a high degree of uncertainty due to the nonequilibrium nature of this plasma. The plasma was over 95 percent ionized at the thruster exit plane. Between 10 and 20 percent of the ions were doubly charged. Two modes of operation were identified. The intensity of plasma emission increased by a factor of two during operation in an oscillatory mode. The transfer between the two modes of operation was likely related to unidentified phenomena occurring on a time scale of minutes.

  18. Thrust Performance of a Low Power Hall Thruster

    NASA Astrophysics Data System (ADS)

    Yamamoto, Naoji; Ezaki, Toru; Nakashima, Hideki

    For the application of Hall thrusters to small satellites, we have been developing a low power Hall thruster. As an initial test, we evaluated the thrust performance of a 50 W class miniature Hall thruster developed at Kyushu University. The outer diameter and the length of the acceleration channel are 18 mm and 7 mm, respectively. A torsional type thrust stand was developed, since the estimated thrust is 1-3 mN, which is too small to be measured by means of a conventional pendulum type thrust stand for Hall thrusts. The uncertainty of this thrust stand is less than 10% at 2 mN with calibration. The thrust, the discharge current, specific impulse and the thrust efficiency at xenon mass flow rate of 0.30 mg/s and discharge voltage of 125 V were 1.7 mN, 0.35 A, 600 sec and 11.4%, respectively. The thrust increased with an increase in discharge voltage and the thrust became 3.1 mN at discharge voltage of 250 V.

  19. High-Efficiency Hall Thruster Discharge Power Converter

    NASA Technical Reports Server (NTRS)

    Jaquish, Thomas

    2015-01-01

    Busek Company, Inc., is designing, building, and testing a new printed circuit board converter. The new converter consists of two series or parallel boards (slices) intended to power a high-voltage Hall accelerator (HiVHAC) thruster or other similarly sized electric propulsion devices. The converter accepts 80- to 160-V input and generates 200- to 700-V isolated output while delivering continually adjustable 300-W to 3.5-kW power. Busek built and demonstrated one board that achieved nearly 94 percent efficiency the first time it was turned on, with projected efficiency exceeding 97 percent following timing software optimization. The board has a projected specific mass of 1.2 kg/kW, achieved through high-frequency switching. In Phase II, Busek optimized to exceed 97 percent efficiency and built a second prototype in a form factor more appropriate for flight. This converter then was integrated with a set of upgraded existing boards for powering magnets and the cathode. The program culminated with integrating the entire power processing unit and testing it on a Busek thruster and on NASA's HiVHAC thruster.

  20. Development of optical diagnostics for performance evaluation of arcjet thrusters

    NASA Technical Reports Server (NTRS)

    Cappelli, Mark A.

    1995-01-01

    Laser and optical emission-based measurements have been developed and implemented for use on low-power hydrogen arcjet thrusters and xenon-propelled electric thrusters. In the case of low power hydrogen arcjets, these laser induce fluorescence measurements constitute the first complete set of data that characterize the velocity and temperature field of such a device. The research performed under the auspices of this NASA grant includes laser-based measurements of atomic hydrogen velocity and translational temperature, ultraviolet absorption measurements of ground state atomic hydrogen, Raman scattering measurements of the electronic ground state of molecular hydrogen, and optical emission based measurements of electronically excited atomic hydrogen, electron number density, and electron temperature. In addition, we have developed a collisional-radiative model of atomic hydrogen for use in conjunction with magnetohydrodynamic models to predict the plasma radiative spectrum, and near-electrode plasma models to better understand current transfer from the electrodes to the plasma. In the final year of the grant, a new program aimed at developing diagnostics for xenon plasma thrusters was initiated, and results on the use of diode lasers for interrogating Hall accelerator plasmas has been presented at recent conferences.

  1. Issues on human acceleration tolerance after long-duration space flights

    NASA Technical Reports Server (NTRS)

    Kumar, K. Vasantha; Norfleet, William T.

    1992-01-01

    This report reviewed the literature on human tolerance to acceleration at 1 G and changes in tolerance after exposure to hypogravic fields. It was found that human tolerance decreased after exposure to hypokinetic and hypogravic fields, but the magnitude of such reduction ranged from 0 to 30 percent for plateau G forces and 30 to 70 percent for time tolerance on sustained G forces. A logistic regression model of the probability of individuals with 25 percent reduction in +Gz tolerance after 1 to 41 days of hypogravic exposures was constructed. The estimated values from the model showed a good correlation with the observed data. A brief review of the need for in-flight centrifuge during long-duration missions was also presented. Review of the available data showed that the use of countermeasures (such as anti-G suits, periodic acceleration, and exercise) reduced the decrement in acceleration tolerance after long-duration space flights. Areas of further research include quantification of the effect of countermeasures on tolerance, and methods to augment tolerance during and after exposures to hypogravic fields. Such data are essential for planning long-duration human missions.

  2. Requirements and Development of an Acceleration Measurement System for International Space Station Microgravity Science Payloads

    NASA Technical Reports Server (NTRS)

    Sutliff, Thomas J.

    1997-01-01

    The International Space Station is being developed by NASA and international partners as a versatile user platform to allow long term on-orbit investigations of a variety of scientific and technology arenas. In particular, scientific studies are planned within a research class known as microgravity science in areas such as biotechnology, combustion, fluid physics, and materials sciences. An acceleration measurement system is in development to aid such research conducted in the on-orbit conditions of apparent weightlessness. This system provides a general purpose acceleration measurement capability in support of these payloads and investigators. Such capability allows for systematic study of scientific phenomena by obtaining information regarding the local accelerations present during experiment operations. Preparations for implementing this flight measurement system involves two distinct stages: requirements development prior to initiating the design activity, and the design activity itself. This paper defines the requirements definition approach taken, provides an overview of the results of the requirements phase, and outlines the initial design considerations being addressed for this measurement system. Some preliminary engineering approaches are also described.

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

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  4. Characterisation of Electrical Propulsion Thrusters as a Load for Electronic Power Supplies

    NASA Astrophysics Data System (ADS)

    Gollor, Matthias; Herty, Frank

    2008-09-01

    For European space missions the importance of electrical propulsion is growing strongly. Many different types of electrical thrusters have been developed in the past years. Often the specification of electrical power supplies suffers from lack of dedicated electrical measurements or misinterpretation of the load behaviour especially under dynamic conditions. As a result, the Power Supply and Control Unit might not be optimally matched to the thrusters and in worst case may react with instabilities during operation.Analyzing the different types of electric thrusters a broad variety of different load characteristics have to be taken into account: non-linear I/V curves, constant voltage and constant current equivalent loads, but except for auxiliary magnet circuits, the loads typically do not show significant inductive or capacitive components. However, the majority of the thrusters show significant load oscillations due to plasma effects, typically in a frequency range of a few ten kHz. Most thrusters are affected by spurious flashovers (sparking, beam-outs, and plasma instabilities).In order to achieve a good definition of the interface between power supplies and the EP thrusters as a load, it is recommended to perform measurements of the current-voltage curve under static and dynamic conditions already in early development phases. For thrusters with complex power supplies the possible coupling between the power sources through the plasma might be considered, too. Examples for such measurements and the transfer of the results into simple electrical models are given for an anode supply of a Kaufmann type ion thruster and a Neutralizer/Keeper supply.

  5. Performance and Thermal Characterization of the NASA-300MS 20 kW Hall Effect Thruster

    NASA Technical Reports Server (NTRS)

    Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Shastry, Rohit; Soulas, George; Smith, Timothy; Mikellides, Ioannis; Hofer, Richard

    2013-01-01

    NASA's Space Technology Mission Directorate is sponsoring the development of a high fidelity 15 kW-class long-life high performance Hall thruster for candidate NASA technology demonstration missions. An essential element of the development process is demonstration that incorporation of magnetic shielding on a 20 kW-class Hall thruster will yield significant improvements in the throughput capability of the thruster without any significant reduction in thruster performance. As such, NASA Glenn Research Center and the Jet Propulsion Laboratory collaborated on modifying the NASA-300M 20 kW Hall thruster to improve its propellant throughput capability. JPL and NASA Glenn researchers performed plasma numerical simulations with JPL's Hall2De and a commercially available magnetic modeling code that indicated significant enhancement in the throughput capability of the NASA-300M can be attained by modifying the thruster's magnetic circuit. This led to modifying the NASA-300M magnetic topology to a magnetically shielded topology. This paper presents performance evaluation results of the two NASA-300M magnetically shielded thruster configurations, designated 300MS and 300MS-2. The 300MS and 300MS-2 were operated at power levels between 2.5 and 20 kW at discharge voltages between 200 and 700 V. Discharge channel deposition from back-sputtered facility wall flux, and plasma potential and electron temperature measurements made on the inner and outer discharge channel surfaces confirmed that magnetic shielding was achieved. Peak total thrust efficiency of 64% and total specific impulse of 3,050 sec were demonstrated with the 300MS-2 at 20 kW. Thermal characterization results indicate that the boron nitride discharge chamber walls temperatures are approximately 100 C lower for the 300MS when compared to the NASA- 300M at the same thruster operating discharge power.

  6. 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.

  7. Performance scaling of gas-fed pulsed plasma thrusters

    NASA Astrophysics Data System (ADS)

    Ziemer, John Kenneth

    The performance scaling of gas-fed pulsed plasma thrusters (GFPPTs) is investigated theoretically and experimentally. Analytical models of the discharge current suggest that close to critically damped current waveforms provide the best energy transfer efficiency. A characteristic velocity for GFPPTs that depends on the inductance-per-unit-length and the square root of the capacitance-to-initial-inductance ratio is also derived in these models. The total efficiency is predicted to be proportional to the ratio of the exhaust velocity to the GFPPT characteristic velocity. A numerical non-dimensional model is used to span a large parameter space of possible operating conditions and suggest optimal configurations. From the non-dimensional model, the exhaust velocity is predicted to scale with a non-dimensional parameter called the dynamic impedance parameter to a power that depends on the mass loading prior to the discharge. To test the validity of the predicted scaling relations, the performance of two rapid-pulse-rate GFPPT designs, PT5 (coaxial electrodes) and PT9 (parallel-plate electrodes), has been measured over 70 different operating conditions with argon propellant. The performance measurements are made in a recently renovated facility that uses liquid nitrogen cooled baffles and a micro-thrust stand capable of measuring impulses <20 muNs within <10%. The measurements demonstrate that the impulse bit scales linearly with the integral of the discharge current squared, as expected for an electromagnetic accelerator. The measured performance scaling in both electrode geometries is shown to be in good agreement with theoretical predictions using the GFPPT characteristic velocity. Normalizing the exhaust velocity and the impulse-to-energy ratio by the GFPPT characteristic velocity collapses almost all the measured data onto single curves that represent the scaling relations for these GFPPTs.

  8. Development of Long-Lifetime Pulsed Gas Valves for Pulsed Electric Thrusters

    NASA Technical Reports Server (NTRS)

    Burkhardt, Wendel M.; Crapuchettes, John M.; Addona, Brad M.; Polzin, Kurt A.

    2015-01-01

    It is advantageous for gas-fed pulsed electric thrusters to employ pulsed valves so propellant is only flowing to the device during operation. The propellant utilization of the thruster will be maximized when all the gas injected into the thruster is acted upon by the fields produced by the electrical pulse. Gas that is injected too early will diffuse away from the thruster before the electrical pulse can act to accelerate the propellant. Gas that is injected too late will miss being accelerated by the already-completed electrical pulse. As a consequence, the valve must open quickly and close equally quickly, only remaining open for a short duration. In addition, the valve must have only a small amount of volume between the sealing body and the thruster so the front and back ends of the pulse are as coincident as possible with the valve cycling, with very little latent propellant remaining in the feed lines after the valve is closed. For a real mission of interest, a pulsed thruster can be expected to pulse at least 10(exp 10) - 10(exp 11) times, setting the range for the number of times a valve must open and close. The valves described in this paper have been fabricated and tested for operation in an inductive pulsed plasma thruster (IPPT) for in-space propulsion. In general, an IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged, producing a high-current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed, it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. The valve characteristics needed for the IPPT application require a fast-acting valve capable of a minimum of 10(exp 10) valve actuation cycles. Since

  9. NASA Brief: Q-Thruster Physics

    NASA Technical Reports Server (NTRS)

    White, Harold

    2013-01-01

    Q-thrusters are a low-TRL form of electric propulsion that operates on the principle of pushing off of the quantum vacuum. A terrestrial analog to this is to consider how a submarine uses its propeller to push a column of water in one direction, while the sub recoils in the other to conserve momentum -the submarine does not carry a "tank" of sea water to be used as propellant. In our case, we use the tools of Magnetohydrodynamics (MHD) to show how the thruster pushes off of the quantum vacuum which can be thought of as a sea of virtual particles -principally electrons and positrons that pop into and out of existence, and where fields are stronger, there are more virtual particles. The idea of pushing off the quantum vacuum has been in the technical literature for a few decades, but to date, the obstacle has been the magnitude of the predicted thrust which has been derived analytically to be very small, and therefore not likely to be useful for human spaceflight. Our recent theoretical model development and test data suggests that we can greatly increase the magnitude of the negative pressure of the quantum vacuum and generate a specific force such that technology based on this approach can be competitive for in-space propulsion approx. 0.1N/kW), and possibly for terrestrial applications (approx. 10N/kW). As an additional validation of the approach, the theory allows calculation of physics constants from first principles: Gravitational constant, Planck constant, Bohr radius, dark energy fraction, electron mass.

  10. Feasibility of a down-scaled HEMP-Thruster as possible N-propulsion system for LISA

    NASA Astrophysics Data System (ADS)

    Keller, A.; Köhler, P.; Gärtner, W.; Hey, F. G.; Berger, M.; Braxmaier, C.; Feili, D.; Weise, D.; Johann, U.

    2013-01-01

    An experimental feasibility study on down-scaling HEMP thrusters to textmu N thrust levels as required e.g. for NGO is presented. Prototypes are used to probe the operation space as well as measuring the divergence angle of the plume and ion acceleration voltage by means of Faraday Cups and a Retarding Potential Analyser in order to gain a deeper understanding of the influence of design parameters. From the measured values thrust and specific impulse are calculated using simple models. Stable operation with a calculated thrust down to 70 N has been demonstrated and divergence efficiencies of about 0.5 are observed. Due to the importance of the thrust value and uncertainties of the models it is clearly desirable to measure the thrust and thrust noise directly with a thrust balance. Such a device is under construction, with a picometer noise level heterodyne interferometer as optical readout.

  11. 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.

  12. Diagnostic evaluations of a beam-shielded 8-cm mercury ion thruster

    NASA Technical Reports Server (NTRS)

    Nakanishi, S.

    1978-01-01

    An engineering model thruster fitted with a remotely actuated graphite fiber polyimide composite beam shield was tested in a 3- by 6.5-meter vacuum facility for in-situ assessment of beam shield effects on thruster performance. Accelerator drain current neutralizer floating potential and ion beam floating potential increased slightly when the shield was moved into position. A target exposed to the low density regions of the ion beam was used to map the boundaries of energetic fringe ions capable of sputtering. The particle efflux was evaluated by measurement of film deposits on cold, heated, bare, and enclosed glass slides.

  13. Electrostatic Plasma Accelerator (EPA)

    NASA Technical Reports Server (NTRS)

    Brophy, John R.; Aston, Graeme

    1995-01-01

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

  14. 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.

  15. Integrated Stirling Convertor and Hall Thruster Test Conducted

    NASA Technical Reports Server (NTRS)

    Mason, Lee S.

    2002-01-01

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

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

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

    NASA's Space Technology Mission Directorate (STMD) Solar Electric Propulsion Technology Demonstration Mission (SEP/TDM) project is funding the development of a 12.5-kW Hall thruster system to support future NASA missions. The thruster designated Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5-kW Hall thruster with magnetic shielding incorporating a centrally mounted cathode. HERMeS was designed and modeled by a NASA GRC and JPL team and was fabricated and tested in vacuum facility 5 (VF5) at NASA GRC. Tests at NASA GRC were performed with the Technology Development Unit 1 (TDU1) thruster. TDU1's magnetic shielding topology was confirmed by measurement of anode potential and low electron temperature along the discharge chamber walls. Thermal characterization tests indicated that during full power thruster operation at peak magnetic field strength, the various thruster component temperatures were below prescribed maximum allowable limits. Performance characterization tests demonstrated the thruster's wide throttling range and found that the thruster can achieve a peak thruster efficiency of 63% at 12.5 kW 500 V and can attain a specific impulse of 3,000 s at 12.5 kW and a discharge voltage of 800 V. Facility background pressure variation tests revealed that the performance, operational characteristics, and magnetic shielding effectiveness of the TDU1 design were mostly insensitive to increases in background pressure.

  17. Development of the Plasma Thruster Particle-in-Cell Simulator to Complement Empirical Studies of a Low-Power Cusped-Field Thruster

    NASA Astrophysics Data System (ADS)

    Gildea, Stephen Robert

    conical jet of fast ions in the near plume region. The influences guiding the formation of the simulated beam in low-current mode are described in detail. A module for predicting erosion rates on dielectric surfaces has also been incorporated into PTpic and applied to simulations of both DCFT operational modes. Two data sets from high---current mode simulations successfully reproduce elevated erosion profiles in each of the three magnetic ring-cusps present in the DCFT. Discrepancies between the simulated and experimental data do exist, however, and are once again attributable to the misplacement of the primary acceleration region of the thruster. Having successfully captured the most significant erosion profile features observed in high-current mode, a simulation of erosion in low-current mode indicates substantially reduced erosion in comparison to the more oscillatory mode. These findings further motivate the completion of low-current mode erosion measurements, and continued numerical studies of the DCFT. Additionally, PTpic has proven to be a useful simulation tool for this project, and has been developed with adaptability in mind to facilitate its application to a variety of thruster designs---including Hall thrusters. (Copies available exclusively from MIT Libraries, libraries.mit.edu/docs - docs mit.edu)

  18. Numerical Modeling and Testing of an Inductively-Driven and High-Energy Pulsed Plasma Thrusters

    NASA Technical Reports Server (NTRS)

    Parma, Brian

    2004-01-01

    Pulsed Plasma Thrusters (PPTs) are advanced electric space propulsion devices that are characterized by simplicity and robustness. They suffer, however, from low thrust efficiencies. This summer, two approaches to improve the thrust efficiency of PPTs will be investigated through both numerical modeling and experimental testing. The first approach, an inductively-driven PPT, uses a double-ignition circuit to fire two PPTs in succession. This effectively changes the PPTs configuration from an LRC circuit to an LR circuit. The LR circuit is expected to provide better impedance matching and improving the efficiency of the energy transfer to the plasma. An added benefit of the LR circuit is an exponential decay of the current, whereas a traditional PPT s under damped LRC circuit experiences the characteristic "ringing" of its current. The exponential decay may provide improved lifetime and sustained electromagnetic acceleration. The second approach, a high-energy PPT, is a traditional PPT with a variable size capacitor bank. This PPT will be simulated and tested at energy levels between 100 and 450 joules in order to investigate the relationship between efficiency and energy level. Arbitrary Coordinate Hydromagnetic (MACH2) code is used. The MACH2 code, designed by the Center for Plasma Theory and Computation at the Air Force Research Laboratory, has been used to gain insight into a variety of plasma problems, including electric plasma thrusters. The goals for this summer include numerical predictions of performance for both the inductively-driven PPT and high-energy PFT, experimental validation of the numerical models, and numerical optimization of the designs. These goals will be met through numerical and experimental investigation of the PPTs current waveforms, mass loss (or ablation), and impulse bit characteristics.

  19. 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.

  20. Effect of Background Pressure on the Performance and Plume of the HiVHAc Hall Thruster

    NASA Technical Reports Server (NTRS)

    Huang, Wensheng; Kamhawi, Hani; Haag, Thomas

    2013-01-01

    During the Single String Integration Test of the NASA HiVHAc Hall thruster, a number of plasma diagnostics were implemented to study the effect of varying facility background pressure on thruster operation. These diagnostics include thrust stand, Faraday probe, ExB probe, and retarding potential analyzer. The test results indicated a rise in thrust and discharge current with background pressure. There was also a decrease in ion energy per charge, an increase in multiply-charged species production, a decrease in plume divergence, and a decrease in ion beam current with increasing background pressure. A simplified ingestion model was applied to determine the maximum acceptable background pressure for thrust measurement. The maximum acceptable ingestion percentage was found to be around 1%. Examination of the diagnostics results suggest the ionization and acceleration zones of the thruster were shifting upstream with increasing background pressure.

  1. Mission Advantages of NEXT: Nasa's Evolutionary Xenon Thruster

    NASA Technical Reports Server (NTRS)

    Oleson, Steven; Gefert, Leon; Benson, Scott; Patterson, Michael; Noca, Muriel; Sims, Jon

    2002-01-01

    With the demonstration of the NSTAR propulsion system on the Deep Space One mission, the range of the Discovery class of NASA missions can now be expanded. NSTAR lacks, however, sufficient performance for many of the more challenging Office of Space Science (OSS) missions. Recent studies have shown that NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system is the best choice for many exciting potential OSS missions including outer planet exploration and inner solar system sample returns. The NEXT system provides the higher power, higher specific impulse, and higher throughput required by these science missions.

  2. Skylab thruster attitude control system

    NASA Technical Reports Server (NTRS)

    Wilmer, G. E., Jr.

    1974-01-01

    Preflight activities and the Skylab mission support effort for the thruster attitude control system (TACS) are documented. The preflight activities include a description of problems and their solutions encountered in the development, qualification, and flight checkout test programs. Mission support effort is presented as it relates to system performance assessment, real-time problem solving, flight anomalies, and the daily system evaluation. Finally, the detailed flight evaluation is presented for each phase of the mission using system telemetry data. Data assert that the TACS met or exceeded design requirements and fulfilled its assigned mission objectives.

  3. Coil system for plasmoid thruster

    NASA Technical Reports Server (NTRS)

    Eskridge, Richard H. (Inventor); Lee, Michael H. (Inventor); Martin, Adam K. (Inventor); Fimognari, Peter J. (Inventor)

    2010-01-01

    A coil system for a plasmoid thruster includes a bias coil, a drive coil and field coils. The bias and drive coils are interleaved with one another as they are helically wound about a conical region. A first field coil defines a first passage at one end of the conical region, and is connected in series with the bias coil. A second field coil defines a second passage at an opposing end of the conical region, and is connected in series with the bias coil.

  4. Current status of MCNP6 as a simulation tool useful for space and accelerator applications

    SciTech Connect

    Mashnik, Stepan G; Bull, Jeffrey S; Hughes, H. Grady; Prael, Richard E; Sierk, Arnold J

    2012-07-20

    For the past several years, a major effort has been undertaken at Los Alamos National Laboratory (LANL) to develop the transport code MCNP6, the latest LANL Monte-Carlo transport code representing a merger and improvement of MCNP5 and MCNPX. We emphasize a description of the latest developments of MCNP6 at higher energies to improve its reliability in calculating rare-isotope production, high-energy cumulative particle production, and a gamut of reactions important for space-radiation shielding, cosmic-ray propagation, and accelerator applications. We present several examples of validation and verification of MCNP6 compared to a wide variety of intermediate- and high-energy experimental data on reactions induced by photons, mesons, nucleons, and nuclei at energies from tens of MeV to about 1 TeV/nucleon, and compare to results from other modern simulation tools.

  5. A direct current rectification scheme for microwave space power conversion using traveling wave electron acceleration

    NASA Technical Reports Server (NTRS)

    Manning, Robert M.

    1993-01-01

    The formation of the Vision-21 conference held three years ago allowed the present author to reflect and speculate on the problem of converting electromagnetic energy to a direct current by essentially reversing the process used in traveling wave tubes that converts energy in the form of a direct current to electromagnetic energy. The idea was to use the electric field of the electromagnetic wave to produce electrons through the field emission process and accelerate these electrons by the same field to produce an electric current across a large potential difference. The acceleration process was that of cyclotron auto-resonance. Since that time, this rather speculative ideas has been developed into a method that shows great promise and for which a patent is pending and a prototype design will be demonstrated in a potential laser power beaming application. From the point of view of the author, a forum such as Vision-21 is becoming an essential component in the rather conservative climate in which our initiatives for space exploration are presently formed. Exchanges such as Vision-21 not only allows us to deviate from the 'by-the-book' approach and rediscover the ability and power in imagination, but provides for the discussion of ideas hitherto considered 'crazy' so that they may be given the change to transcend from the level of eccentricity to applicability.

  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. A torsion balance for impulse and thrust measurements of micro-Newton thrusters.

    PubMed

    Yang, Yuan-Xia; Tu, Liang-Cheng; Yang, Shan-Qing; Luo, Jun

    2012-01-01

    This paper reports the performance of a torsion-type thrust stand suitable for studies of micro-Newton thrusters, which is developed for ground testing the micro-Newton thruster in Chinese Test of the Equivalence Principle with Optical readout space mission. By virtue of specially suspending design and precise assembly of torsion balance configuration, the thrust stand with load capacity up to several kilograms is able to measure the impulse bit up to 1350 μNs with a resolution of 0.47 μNs, and the average thrust up to 264 μN with a resolution of 0.09 μN in both open and close loop operation. A pulsed plasma thruster, the preliminary prototype developed for Chinese TEPO space mission, is tested by the thrust stand, and the results reveal that the average impulse bit per pulse is measured to be 58.4 μNs with a repeatability of about 5%.

  8. 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.

  9. 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.

  10. Mode transition of a Hall thruster discharge plasma

    SciTech Connect

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

    2014-05-28

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

  11. NASA's Evolutionary Xenon Thruster (NEXT) Component Verification Testing

    NASA Technical Reports Server (NTRS)

    Herman, Daniel A.; Pinero, Luis R.; Sovey, James S.

    2009-01-01

    Component testing is a critical facet of the comprehensive thruster life validation strategy devised by the NASA s Evolutionary Xenon Thruster (NEXT) program. Component testing to-date has consisted of long-duration high voltage propellant isolator and high-cycle heater life validation testing. The high voltage propellant isolator, a heritage design, will be operated under different environmental condition in the NEXT ion thruster requiring verification testing. The life test of two NEXT isolators was initiated with comparable voltage and pressure conditions with a higher temperature than measured for the NEXT prototype-model thruster. To date the NEXT isolators have accumulated 18,300 h of operation. Measurements indicate a negligible increase in leakage current over the testing duration to date. NEXT 1/2 in. heaters, whose manufacturing and control processes have heritage, were selected for verification testing based upon the change in physical dimensions resulting in a higher operating voltage as well as potential differences in thermal environment. The heater fabrication processes, developed for the International Space Station (ISS) plasma contactor hollow cathode assembly, were utilized with modification of heater dimensions to accommodate a larger cathode. Cyclic testing of five 1/22 in. diameter heaters was initiated to validate these modified fabrication processes while retaining high reliability heaters. To date two of the heaters have been cycled to 10,000 cycles and suspended to preserve hardware. Three of the heaters have been cycled to failure giving a B10 life of 12,615 cycles, approximately 6,000 more cycles than the established qualification B10 life of the ISS plasma contactor heaters.

  12. Motion-Based System Identification and Fault Detection and Isolation Technologies for Thruster Controlled Spacecraft

    NASA Technical Reports Server (NTRS)

    Wilson, Edward; Sutter, David W.; Berkovitz, Dustin; Betts, Bradley J.; Kong, Edmund; delMundo, Rommel; Lages, Christopher R.; Mah, Robert W.; Papasin, Richard

    2003-01-01

    By analyzing the motions of a thruster-controlled spacecraft, it is possible to provide on-line (1) thruster fault detection and isolation (FDI), and (2) vehicle mass- and thruster-property identification (ID). Technologies developed recently at NASA Ames have significantly improved the speed and accuracy of these ID and FDI capabilities, making them feasible for application to a broad class of spacecraft. Since these technologies use existing sensors, the improved system robustness and performance that comes with the thruster fault tolerance and system ID can be achieved through a software-only implementation. This contrasts with the added cost, mass, and hardware complexity commonly required by FDI. Originally developed in partnership with NASA - Johnson Space Center to provide thruster FDI capability for the X-38 during re-entry, these technologies are most recently being applied to the MIT SPHERES experimental spacecraft to fly on the International Space Station in 2004. The model-based FDI uses a maximum-likelihood calculation at its core, while the ID is based upon recursive least squares estimation. Flight test results from the SPHERES implementation, as flown aboard the NASA KC-1 35A 0-g simulator aircraft in November 2003 are presented.

  13. In-Space Propulsion Technology Program Solar Electric Propulsion Technologies

    NASA Technical Reports Server (NTRS)

    Dankanich, John W.

    2006-01-01

    NASA's In-space Propulsion (ISP) Technology Project is developing new propulsion technologies that can enable or enhance near and mid-term NASA science missions. The Solar Electric Propulsion (SEP) technology area has been investing in NASA s Evolutionary Xenon Thruster (NEXT), the High Voltage Hall Accelerator (HiVHAC), lightweight reliable feed systems, wear testing, and thruster modeling. These investments are specifically targeted to increase planetary science payload capability, expand the envelope of planetary science destinations, and significantly reduce the travel times, risk, and cost of NASA planetary science missions. Status and expected capabilities of the SEP technologies are reviewed in this presentation. The SEP technology area supports numerous mission studies and architecture analyses to determine which investments will give the greatest benefit to science missions. Both the NEXT and HiVHAC thrusters have modified their nominal throttle tables to better utilize diminished solar array power on outbound missions. A new life extension mechanism has been implemented on HiVHAC to increase the throughput capability on low-power systems to meet the needs of cost-capped missions. Lower complexity, more reliable feed system components common to all electric propulsion (EP) systems are being developed. ISP has also leveraged commercial investments to further validate new ion and hall thruster technologies and to potentially lower EP mission costs.

  14. Investigation of a repetitive pulsed electrothermal thruster

    NASA Technical Reports Server (NTRS)

    Burton, R. L.; Fleischer, D.; Goldstein, S. A.; Tidman, D. A.; Winsor, N. K.

    1986-01-01

    A pulsed electrothermal (PET) thruster with 1000:1 ratio nozzle is tested in a repetitive mode on water propellant. The thruster is driven by a 60J pulse forming network at repetition rates up to 10 Hz (600W). The pulse forming network has a .31 ohm impedance, well matched to the capillary discharge resistance of .40 ohm, and is directly coupled to the thruster electrodes without a switch. The discharge is initiated by high voltage breakdown, typically at 2500V, through the water vapor in the interelectrode gap. Water is injected as a jet through a .37 mm orifice on the thruster axis. Thruster voltage, current and impulse bit are recorded for several seconds at various power supply currents. Thruster to power ratio is typically T/P = .07 N/kW. Tank background pressure precludes direct measurement of exhaust velocity which is inferred from calculated pressure and temperature in the discharge to be about 14 km/sec. Efficiency, based on this velocity and measured T/P is .54 + or - .07. Thruster ablation is zero at the throat and becomes measurable further upstream, indicating that radiative ablation is occurring late in the pulse.

  15. Progress on the Plasmoid Thruster Experiment (PTX)

    NASA Technical Reports Server (NTRS)

    Martin, Adam; Eskridge, Richard; Fimognari, Peter; Koelfgen, Syri J.; Lee, Mike

    2004-01-01

    A plasmoid is a compact plasma structure with an integral magnetic field, that may be categorized according to the relative strength of the poloidal and toroidal magnetic field (B(sub p) and B(sub t), respectively). An object with B(sub p)/B(sub t), much much more than 1 is called a Field Reverse Configuration (FRC); if B(sub p) approximately equal to B(sub t), it is called a Spheromak. The thruster operates by repetitively producing plasmoids that are accelerated and ejected at high velocity. As this process is inductive, there are no electrodes. Also, the magnetic structure of the plasmoid should suppress thermal and mass losses to the wall, and improve detachment of the plasma exhaust from the thruster. This concept should be capable of producing an Isp in the range of 5,000 - 10,000 s with thrust densities of order 10(exp 5) N per square meters. The current experiment is designed to produce jet powers in the range of 5-10 kW, although the concept should be scalable into the MW range. In PTX, the plasmoid is formed inside of a single turn conical theta-pinch coil (17.58 cone angle). The coil is driven by a 640 nF, 35 kV capacitor bank, which rings at a frequency of 500 kHz. Previous experiments on PTX were conducted with a static-fill of propellant gas (6% H2 in He), and demonstrated reliable ionization over a pressure range of 40 - 200 mTorr. We are now adding a fast gas-puff valve to load the propellant, and a ringing pre-ionization circuit (f = 5 Mhz) to better control the plasmoid formation. An alternate coil (8.58 cone angle) will also be used, so as to investigate the effect of coil shape on performance. In addition, a variety of propellants will be used, including hydrogen, nitrogen, and argon. The plasmoid mass and velocity will be measured with a variety of diagnostics, including external B-dot probes and flux loops, a high-speed framing camera, and a HeNe laser interferometer. Internal B-dot probes and a quadruple Langmuir probe will provide additional

  16. Single bunch transverse instability in a circular accelerator with chromaticity and space charge

    SciTech Connect

    Balbekov, V.

    2015-10-21

    The transverse instability of a bunch in a circular accelerator is elaborated in this paper. A new tree-modes model is proposed and developed to describe the most unstable modes of the bunch. This simple and flexible model includes chromaticity and space charge, and can be used with any bunch and wake forms. The dispersion equation for the bunch eigentunes is obtained in form of a third-order algebraic equation. The known head-tail and TMCI modes appear as the limiting cases which are distinctly bounded at zero chromaticity only. It is shown that the instability parameters depend only slightly on the bunch model but they are rather sensitive to the wake shape. In particular, space charge effects are investigated in the paper and it is shown that their influence depends on sign of wake field enhancing the bunch stability if the wake is negative. In addition, the resistive wall wake is considered in detail including a comparison of single and collective effects. A comparison of the results with earlier publications is carried out.

  17. Single bunch transverse instability in a circular accelerator with chromaticity and space charge

    DOE PAGESBeta

    Balbekov, V.

    2015-10-21

    The transverse instability of a bunch in a circular accelerator is elaborated in this paper. A new tree-modes model is proposed and developed to describe the most unstable modes of the bunch. This simple and flexible model includes chromaticity and space charge, and can be used with any bunch and wake forms. The dispersion equation for the bunch eigentunes is obtained in form of a third-order algebraic equation. The known head-tail and TMCI modes appear as the limiting cases which are distinctly bounded at zero chromaticity only. It is shown that the instability parameters depend only slightly on the bunchmore » model but they are rather sensitive to the wake shape. In particular, space charge effects are investigated in the paper and it is shown that their influence depends on sign of wake field enhancing the bunch stability if the wake is negative. In addition, the resistive wall wake is considered in detail including a comparison of single and collective effects. A comparison of the results with earlier publications is carried out.« less

  18. 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.

  19. Ion thruster charge-exchange plasma flow

    NASA Technical Reports Server (NTRS)

    Carruth, M. R., Jr.; Gabriel, S. B.; Kitamura, S.

    1982-01-01

    The electron bombardment ion thruster has been under development for a number of years and during this time, studies of the plasmas produced by the thrusters and their interactions with spacecraft have been evaluated, based on available data. Due to diagnostic techniques used and facility effects, there is uncertainty as to the reliability of data from these early studies. This paper presents data on the flow of the charge-exchange plasma produced just downstream of the thruster's ion optics. The 'end-effect' of a cylindrical Langmuir probe is used to determine ion density and directed ion velocity. Results are compared with data obtained from a retarding potential analyzer-Faraday cup.

  20. Ion thruster plume effects on spacecraft surfaces

    NASA Technical Reports Server (NTRS)

    Carruth, M. R., Jr.; Kuo, Y. S.

    1981-01-01

    A charge-exchange plasma, generated by an ion thruster, is capable of flowing upstream from the ion thruster and therefore represents a source of contamination to a spacecraft. An analytical model of the charge-exchange plasma density around a spacecraft was used to estimate the contamination which various spacecraft materials may be exposed to. Measurements of plasma density around an ion thruster were compared to this model. Results of experimental studied regarding the effects on various spacecraft materials' properties due to exposure to expected mercury contamination levels are presented.

  1. Failure Investigation of an Intra-Manifold Explosion in a Horizontally-Mounted 870 lbf Reaction Control Thruster

    NASA Technical Reports Server (NTRS)

    Durning, Joseph G., III; Westover, Shayne C.; Cone, Darren M.

    2011-01-01

    In June 2010, an 870 lbf Space Shuttle Orbiter Reaction Control System Primary Thruster experienced an unintended shutdown during a test being performed at the NASA White Sands Test Facility. Subsequent removal and inspection of the thruster revealed permanent deformation and misalignment of the thruster valve mounting plate. Destructive evaluation determined that after three nominal firing sequences, the thruster had experienced an energetic event within the fuel (monomethylhydrazine) manifold at the start of the fourth firing sequence. The current understanding of the phenomenon of intra-manifold explosions in hypergolic bipropellant thrusters is documented in literature where it is colloquially referred to as a ZOT. The typical ZOT scenario involves operation of a thruster in a gravitational field with environmental pressures above the triple point pressure of the propellants. Post-firing, when the thruster valves are commanded closed, there remains a residual quantity of propellant in both the fuel and oxidizer (nitrogen tetroxide) injector manifolds known as the "dribble volume". In an ambient ground test configuration, these propellant volumes will drain from the injector manifolds but are impeded by the local atmospheric pressure. The evacuation of propellants from the thruster injector manifolds relies on the fluids vapor pressure to expel the liquid. The higher vapor pressure oxidizer will evacuate from the manifold before the lower vapor pressure fuel. The localized cooling resulting from the oxidizer boiling during manifold draining can result in fuel vapor migration and condensation in the oxidizer passage. The liquid fuel will then react with the oxidizer that enters the manifold during the next firing and may produce a localized high pressure reaction or explosion within the confines of the oxidizer injector manifold. The typical ZOT scenario was considered during this failure investigation, but was ultimately ruled out as a cause of the explosion

  2. Where is the breathing mode? High voltage Hall effect thruster studies with EMD method

    SciTech Connect

    Kurzyna, J.; Makowski, K.; Mazouffre, S.; Peradzynski, Z.; Lazurenko, A.; Dudeck, M.

    2008-03-19

    Discharge current and local plasma oscillations are studied in a high voltage Hall effect thruster PPS registered -X000. Characteristic time scales that appear in different operating conditions are resolved with the use of Hilbert-Huang spectra (HHS) which display time dependenc of instantaneous frequency and power. Sets of intrinsic mode functions (imfs) that are used for HHS calculation result due to application of empirical mode decomposition method (EMD) to nonstationary multicomponent signals. In the experiment the signals are captured in the electric circuit of the thruster as well locally, in the vicinity of the thruster exhaust region. Classical electric probes spaced along the azimuth and/or thruster axis let us study correlations of signals which were captured in different locations. In this way azimuthal and axial propagation of disturbances is inspected. The discharge voltage is varied in the range of 200 divide 900 V while xenon mas flow rate of 5 divide 9 mg/s. LF, MF, and HF characteristic bands that are known from previous studies of PPS registered -100 thruster have been also detected here. However, expanding discharge current onto a set of intrinsic modes we can resolve MF mode more reliably than before. Moreover, for higher discharge voltages, this irregular mode turns into more regular waveform and tends to dominate in the discharge current masking almost completely the breathing mode (LF oscillations of the discharge current). In such a case triggering of HF oscillations is correlated with the phase of MF mode while in the case of PPS registered -100 thruster it was correlated with the appropriate phase of the breathing mode (LF band). Regular HF emission that can be unambiguously interpreted as azimuthal electrostatic wave now is observed only in the specific operating conditions of the thruster. However, even if irregular HF emission is observed the time delay of cross-correlated signals which are captured in different azimuthal locations

  3. Performance of a Cylindrical Hall-Effect Thruster Using Permanent Magnets

    NASA Technical Reports Server (NTRS)

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

    2009-01-01

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

  4. Lunar Reconnaissance Orbiter (LRO) Thruster Control Mode Design and Flight Experience

    NASA Technical Reports Server (NTRS)

    Hsu, Oscar C.

    2010-01-01

    National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, designed, built, tested, and launched the Lunar Reconnaissance Orbiter (LRO) from Cape Canaveral Air Force Station on June 18, 2009. The LRO spacecraft is the first operational spacecraft designed to support NASA s return to the Moon, as part of the Vision for Space Exploration. LRO was launched aboard an Atlas V 401 launch vehicle into a direct insertion trajectory to the Moon. Twenty-four hours after separation the propulsion system was used to perform a mid-course correction maneuver. Four days after the mid-course correction a series of propulsion maneuvers were executed to insert LRO into its commissioning orbit. The commission period lasted eighty days and this followed by a second set of thruster maneuvers that inserted LRO into its mission orbit. To date, the spacecraft has been gathering invaluable data in support of human s future return to the moon. The LRO Attitude Control Systems (ACS) contains two thruster based control modes: Delta-H and Delta-V. The design of the two controllers are similar in that they are both used for 3-axis control of the spacecraft with the Delta-H controller used for momentum management and the Delta-V controller used for orbit adjust and maintenance maneuvers. In addition to the nominal purpose of the thruster modes, the Delta-H controller also has the added capability of performing a large angle slew maneuver. A suite of ACS components are used by the thruster based control modes, for both initialization and control. For initialization purposes, a star tracker or the Kalman Filter solution is used for providing attitude knowledge and upon entrance into the thruster based control modes attitude knowledge is provided via rate propagation using a inertial reference unit (IRU). Rate information for the controller is also supplied by the IRU. Three-axis control of the spacecraft in the thruster modes is provided by eight 5

  5. ATS-F radiant cooler contamination test in a hydrazine thruster exhaust

    NASA Technical Reports Server (NTRS)

    Chirivella, J. E.

    1973-01-01

    A test was conducted under simulated space conditions to determine the potential thermal degradation of the ATS-F radiant cooler from any contaminants generated by a 0.44-N(0.1-lbf) hydrazine thruster. The radiant cooler, a 0.44-N(0.1-lbf)hydrazine engine, and an aluminum plate simulating the satellite interface were assembled to simulate their flight configuration. The cooler was provided with platinum sensors for measuring temperature, and its surfaces were instrumented with six quartz crystal microbalance units (QCM) to measure contaminant mass deposits. The complete assembly was tested in the molecular sink vacuum facility (Molsink) at the Jet Propulsion Laboratory. This was the first time that a radiant cooler and a hydrazine engine were tested together in a very-high-vacuum space simulator, and this test was the first successful measurement of detectable deposits from hydrazine rocket engine plumes in a high vacuum. The engine was subjected to an accelerated duty cycle of 1 pulse/min, and after 2-hr of operation, the QCMs began to shift in frequency. The tests continued for several days and, although there was considerable activity in the QCMs, the cooler never experienced thermal degradation.

  6. Status of the NEXT Ion Thruster Long Duration Test

    NASA Technical Reports Server (NTRS)

    Frandina, Michael M.; Arrington, Lynn A.; Soulas, George C.; Hickman, Tyler A.; Patterson, Michael J.

    2005-01-01

    The status of NASA's Evolutionary Xenon Thruster (NEXT) Long Duration Test (LDT) is presented. The test will be conducted with a 36 cm diameter engineering model ion thruster, designated EM3, to validate and qualify the NEXT thruster propellant throughput capability of 450 kg xenon. The ion thruster will be operated at various input powers from the NEXT throttle table. Pretest performance assessments demonstrated that EM3 satisfies all thruster performance requirements. As of June 26, 2005, the ion thruster has accumulated 493 hours of operation and processed 10.2 kg of xenon at a thruster input power of 6.9 kW. Overall ion thruster performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, has been steady to date with very little variation in performance parameters.

  7. 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

  8. Accelerator-Based Studies of Heavy Ion Interactions Relevant to Space Biomedicine

    NASA Technical Reports Server (NTRS)

    Miller, J.; Heilbronn, L.; Zeitlin, C.

    1999-01-01

    Evaluation of the effects of space radiation on the crews of long duration space missions must take into account the interactions of high energy atomic nuclei in spacecraft and planetary habitat shielding and in the bodies of the astronauts. These heavy ions (i.e. heavier than hydrogen), while relatively small in number compared to the total galactic cosmic ray (GCR) charged particle flux, can produce disproportionately large effects by virtue of their high local energy deposition: a single traversal by a heavy charged particle can kill or, what may be worse, severely damage a cell. Research into the pertinent physics and biology of heavy ion interactions has consequently been assigned a high priority in a recent report by a task group of the National Research Council. Fragmentation of the incident heavy ions in shielding or in the human body will modify an initially well known radiation field and thereby complicate both spacecraft shielding design and the evaluation of potential radiation hazards. Since it is impractical to empirically test the radiation transport properties of each possible shielding material and configuration, a great deal of effort is going into the development of models of charged particle fragmentation and transport. Accurate nuclear fragmentation cross sections (probabilities), either in the form of measurements with thin targets or theoretical calculations, are needed for input to the transport models, and fluence measurements (numbers of fragments produced by interactions in thick targets) are needed both to validate the models and to test specific shielding materials and designs. Fluence data are also needed to characterize the incident radiation field in accelerator radiobiology experiments. For a number of years, nuclear fragmentation measurements at GCR-like energies have been carried out at heavy ion accelerators including the LBL Bevalac, Saturne (France), the Synchrophasotron and Nuklotron (Dubna, Russia), SIS-18 (GSI, Germany), the

  9. Thruster models for NEP system analysis

    NASA Technical Reports Server (NTRS)

    Gilland, Jim

    1993-01-01

    There are currently no thruster modeling codes that can be integrated with power system codes for full propulsion system modeling. Most existing thruster models were written from a 'stand alone' viewpoint, assuming the user is performing analyses on thruster performance alone. The goal of the present modeling effort is to develop thruster codes that model performance and scaling as a function of mission and system inputs, rather than in terms of more elemental physical parameters. System level parameters of interest are as follows: performance, such as specific impulse and efficency; terminal characteristics, such as voltage or current; and mass. Specific impulse and efficiency couple with mission analyses, while terminal characteristics allow integration with power systems. Additional information on lifetime and operation may be required for detailed designs.

  10. The MPD thruster program at JPL

    NASA Astrophysics Data System (ADS)

    Goodfellow, Keith; Pivirotto, Tom; Polk, Jay

    The topics covered are presented in viewgraph form and include the following: engine lifetime assessment; lithium magnetoplasmadynamic (MPD) thruster development; and radiation-cooled, applied-field engine testing.

  11. The NASA GSFC MEMS Colloidal Thruster

    NASA Technical Reports Server (NTRS)

    Cardiff, Eric H.; Jamieson, Brian G.; Norgaard, Peter C.; Chepko, Ariane B.

    2004-01-01

    A number of upcoming missions require different thrust levels on the same spacecraft. A highly scaleable and efficient propulsion system would allow substantial mass savings. One type of thruster that can throttle from high to low thrust while maintaining a high specific impulse is a Micro-Electro-Mechanical System (MEMS) colloidal thruster. The NASA GSFC MEMS colloidal thruster has solved the problem of electrical breakdown to permit the integration of the electrode on top of the emitter by a novel MEMS fabrication technique. Devices have been successfully fabricated and the insulation properties have been tested to show they can support the required electric field. A computational finite element model was created and used to verify the voltage required to successfully operate the thruster. An experimental setup has been prepared to test the devices with both optical and Time-Of-Flight diagnostics.

  12. 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.

  13. 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.

  14. Study of monopropellants for electrothermal thrusters

    NASA Technical Reports Server (NTRS)

    Kuenzly, J. D.

    1974-01-01

    A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.

  15. An acceleration of the characteristics by a space-angle two-level method using surface discontinuity factors

    SciTech Connect

    Grassi, G.

    2006-07-01

    We present a non-linear space-angle two-level acceleration scheme for the method of the characteristics (MOC). To the fine level on which the MOC transport calculation is performed, we associate a more coarsely discretized phase space in which a low-order problem is solved as an acceleration step. Cross sections on the coarse level are obtained by a flux-volume homogenisation technique, which entails the non-linearity of the acceleration. Discontinuity factors per surface are introduced as additional degrees of freedom on the coarse level in order to ensure the equivalence of the heterogeneous and the homogenised problem. After each fine transport iteration, a low-order transport problem is iteratively solved on the homogenised grid. The solution of this problem is then used to correct the angular moments of the flux resulting from the previous free transport sweep. Numerical tests for a given benchmark have been performed. Results are discussed. (authors)

  16. Detailed analysis of Honeywell In-Space Accelerometer data - STS-32. [crystal microstructure response to different types of residual acceleration

    NASA Technical Reports Server (NTRS)

    Rogers, Melissa J. B.; Alexander, J. I. D.; Schoess, Jeff

    1993-01-01

    The Honeywell In-Space Accelerometer (HISA) system collected data in the mid-deck area of the Shuttle Columbia during the flight of STS-32, January 1990. The resulting data were to be used to investigate the response of crystal microstructure to different types of residual acceleration. The HISA is designed to detect and record transient and oscillatory accelerations. The sampling and electronics package stored averaged accelerations over two sampling periods; two sampling rates were available: 1 Hz and 50 Hz. Analysis of the HISA data followed the CMMR Acceleration Data Processing Guide, considering in-house computer modelling of a float-zone indium crystal growth experiment. Characteristic examples of HISA data showing the response to the primary reaction control system, Orbiter Maneuvering System operations, and crew treadmill activity are presented. Various orbiter structural modes are excited by these and other activities.

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

    NASA Technical Reports Server (NTRS)

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

    2013-01-01

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

  18. The Green Propellant Infusion Mission Thruster Performance Testing for Plume Diagnostics

    NASA Technical Reports Server (NTRS)

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

    2014-01-01

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

  19. 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.

  20. 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.

  1. The 6670-Newton attitude-control thruster using hydrogen-oxygen propellant

    NASA Technical Reports Server (NTRS)

    Gordon, L. H.

    1977-01-01

    The development of a reusable, attitude-control propulsion system for the space transportation system is discussed. A flight weight, gaseous oxygen attitude control thruster assembly was tested to obtain data on cyclic life, thermal and hydraulic characteristics, pulse response, and performance. The basic thruster components were tested in excess of 51,000 pulses and 660 seconds, steady state, with no degradation of the 93 percent characteristic exhaust velocity efficiency level. Nominal operating conditions were a chamber pressure of 207 N sq cm (300 psia), a mixture ratio of 4.0, a pulse width of 100 ms, and a pulse frequency of 2 Hz.

  2. Liquid-Metal-Fed Pulsed Plasma Thrusters

    NASA Technical Reports Server (NTRS)

    Markusic, Thomas

    2003-01-01

    Liquid metal propellants may provide a path toward more reliable and efficient pulsed plasma thrusters (PPTs). Conceptual thruster designs which eliminate the need for high current switches and propellant metering valves are described. Propellant loading techniques are suggested that, at least in principle, show promise to increase propellant utilization, dynamic, and electrical efficiency. Experimental results from a prototype electromagnetically-pumped propellant feed system, and experiments in the initiation of arc discharges in liquid metal droplets, are presented.

  3. MPD thruster research issues, activities, strategies

    NASA Technical Reports Server (NTRS)

    1991-01-01

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

  4. Electrothermal thruster diagnostics. Volume 2: Technical

    NASA Technical Reports Server (NTRS)

    Zafran, S.; Jackson, B.

    1983-01-01

    Test data taken with nitrogen, hydrogen, and ammonia propellants are presented over a wide range of operating conditions at thrust levels up to 225 mN (50 mlb). The design adaptation of a flight-qualified thruster for operation with gaseous propellant inlet is described. Post-test analysis includes evaluation of thruster performance and efficiency, and shows the effects of propellant contamination on an immersed heating element.

  5. Electrothermal thruster diagnostics. Volume 1: Executive summary

    NASA Technical Reports Server (NTRS)

    Zafran, S.; Jackson, B.

    1983-01-01

    A flight-qualified electrothermal thruster demonstrated its adaptability to a variety of propellants. Originally qualified for operation with hydrazine propellant, it was operated with nitrogen, hydrogen, and ammonia propellants, demonstrating 73, 61, and 52 percent overall efficiency with these propellants, respectively, when tested over a wide range of operating conditions. By introducing a preheater to admit hot, rather than cold, propellant inlet gases to the thruster's augmentation heat exchanger, delivered specific impulse closer to theoretical performance limits should be achieved.

  6. Pulsed Plasma Thruster Technology Development and Flight Demonstration Program

    NASA Technical Reports Server (NTRS)

    Peterson, Todd T.; Curran, Frank M.

    1998-01-01

    Because of anticipated near- and far-term mission needs for innovative new electric propulsion technologies, the NASA On-Board Propulsion program is sponsoring NASA Lewis Research Centers' Pulsed Plasma Thruster (PPT) technology development program to rapidly advance PPT technology. This is a joint effort of Lewis' On-Board Propulsion Branch and Space Flight Project Branch. Building on state-of-the-art technology developed and flown in the 1970's and 1980's for Department of Defense missions, newly developed PPT program designs are realizing significant gains in system mass reduction, thrust to mass ratio, total impulse to mass ratio, and variability of impulse bit.

  7. Performance of a magnetic multipole line-cusp argon ion thruster

    NASA Technical Reports Server (NTRS)

    Sovey, J. S.

    1981-01-01

    A 17 cm diameter line cusp ion thruster was evaluated with inert gases which are candidate propellants for on orbit and orbit transfer propulsion functions for Large Space Systems. A semiempirical relationship was generated to predict thruster beam current in terms of plasma parameters which would allow initial thruster optimization without ion extraction and the associated large vacuum facilities. The sensitivity of performance to changes in discharge electrode configurations and magnetic circuit was evaluated and is presented. After final optimization a propellant utilization efficiency of 0.9 at a discharge chamber power expenditure of about 260 w per beam ampere was obtained. These performance parameters are the highest yet achieved with argon propellant.

  8. Extended performance solar electric propulsion thrust system study. Volume 4: Thruster technology evaluation

    NASA Technical Reports Server (NTRS)

    Poeschel, R. L.; Hawthorne, E. I.; Weisman, Y. C.; Frisman, M.; Benson, G. C.; Mcgrath, R. J.; Martinelli, R. M.; Linsenbardt, T. L.; Beattie, J. R.

    1977-01-01

    Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. Emphasis was placed on relatively high power missions (60 to 100 kW) such as a Halley's comet rendezvous. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed in sufficient detail for comparing mass, efficiency, reliability, structure, and thermal characteristics. Confirmation testing and analysis of thruster and power processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. The baseline thrust system design features modular construction, conventional power processing, and a concentrator solar array concept and is designed to interface with the Space Shuttle.

  9. 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.

  10. Low power arcjet thruster pulse ignition

    NASA Technical Reports Server (NTRS)

    Sarmiento, Charles J.; Gruber, Robert P.

    1987-01-01

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

  11. Low Frequency Vibration Characteristics of the Space Acceleration Measurement System 2 Tape Drive Assembly

    NASA Technical Reports Server (NTRS)

    Javeed, Mehzad; Russell, James W.

    1996-01-01

    This report summarizes results of force and moment measurements of the Space Acceleration Measurement System 2 (SAMS 2) Tape Drive Assembly (TDA) over the frequency range from 0.35 Hz to 256 Hz for steady state operations including write, read, rewind, and fast forward. Time domain force results are presented for transient TDA operations that include software eject, manual eject, and manual load. Three different mounting configurations were employed for attaching the inner box with the tape drive unit to the outer box. Two configurations employed grommet sets with spring rates of 42 and 62 pounds per inch respectively. The third configuration employed a set of metallic washers. For all four steady state operations the largest average forces were on the Y axis with the metallic washers and were less than 0.005 pounds. The largest average moments were on the X axes with the washers and were less than 0.030 pound inches. At the third octave centerband frequency of 31.5 Hz, the 42 pound per inch grommets showed the greatest forces and moments for read and write operations. At the third octave centerband frequency of 49.6 Hz, the 62 pound per inch grommets showed the greatest forces and moments for rewind operation. Transient operation forces ranged from 0.75 pounds for the software eject to greater than 1 pound for manual load and eject.

  12. Iodine Hall Thruster Propellant Feed System for a CubeSat

    NASA Technical Reports Server (NTRS)

    Polzin, Kurt A.

    2014-01-01

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

  13. Reduced Toxicity Fuel Satellite Propulsion System Including Axial Thruster and ACS Thruster Combination

    NASA Technical Reports Server (NTRS)

    Schneider, Steven J. (Inventor)

    2001-01-01

    A reduced toxicity fuel satellite propulsion system including a reduced toxicity propellant supply for consumption in an axial class thruster and an ACS class thruster. The system includes suitable valves and conduits for supplying the reduced toxicity propellant to the ACS decomposing element of an ACS thruster. The ACS decomposing element is operative to decompose the reduced toxicity propellant into hot propulsive gases. In addition the system includes suitable valves and conduits for supplying the reduced toxicity propellant to an axial decomposing element of the axial thruster. The axial decomposing element is operative to decompose the reduced toxicity propellant into hot gases. The system further includes suitable valves and conduits for supplying a second propellant to a combustion chamber of the axial thruster, whereby the hot gases and the second propellant auto-ignite and begin the combustion process for producing thrust.

  14. 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

  15. Space-charge effects in ultra-high current electron bunches generated by laser-plasma accelerators

    SciTech Connect

    Grinner, F. J.; Schroeder, C. B.; Maier, A. R.; Becker, S.; Mikhailova, J. M.

    2009-02-11

    Recent advances in laser-plasma accelerators, including the generation of GeV-scale electron bunches, enable applications such as driving a compact free-electron-laser (FEL). Significant reduction in size of the FEL is facilitated by the expected ultra-high peak beam currents (10-100 kA) generated in laser-plasma accelerators. At low electron energies such peak currents are expected to cause space-charge effects such as bunch expansion and induced energy variations along the bunch, potentially hindering the FEL process. In this paper we discuss a self-consistent approach to modeling space-charge effects for the regime of laser-plasma-accelerated ultra-compact electron bunches at low or moderate energies. Analytical treatments are considered as well as point-to-point particle simulations, including the beam transport from the laser-plasma accelerator through focusing devices and the undulator. In contradiction to non-self-consistent analyses (i.e., neglecting bunch evolution), which predict a linearly growing energy chirp, we have found the energy chirp reaches a maximum and decreases thereafter. The impact of the space-charge induced chirp on FEL performance is discussed and possible solutions are presented.

  16. Effects of Hyperbolic Rotation in Minkowski Space on the Modeling of Plasma Accelerators in a Lorentz Boosted Frame

    SciTech Connect

    Vay, J.-L.; Geddes, C. G. R.; Cormier-Michel, E.; Grote, D. P.

    2010-09-21

    Laser driven plasma accelerators promise much shorter particle accelerators but their development requires detailed simulations that challenge or exceed current capabilities. We report the first direct simulations of stages up to 1 TeV from simulations using a Lorentz boosted calculation frame resulting in a million times speedup, thanks to a frame boost as high as gamma = 1300. Effects of the hyperbolic rotation in Minkowski space resulting from the frame boost on the laser propagation in the plasma is shown to be key in the mitigation of a numerical instability that was limiting previous attempts.

  17. DC-like Phase Space Manipulation and Particle Acceleration Using Chirped AC Fields

    SciTech Connect

    P.F. Schmit and N.J. Fisch

    2009-06-17

    Waves in plasmas can accelerate particles that are resonant with the wave. A DC electric field also accelerates particles, but without a resonance discrimination, which makes the acceleration mechanism profoundly different. We investigate the effect on a Hamiltonian distribution of an accelerating potential waveform, which could, for example, represent the average ponderomotive effect of two counterpropagating electromagnetic waves. In particular, we examine the apparent DC-like time-asymptotic response of the distribution in regimes where the potential structure is accelerated adiabatically. A highly resonant population within the distribution is always present, and we characterize its nonadiabatic response during wave-particle resonance using an integral method in the noninertial reference frame moving with the wave. Finally, we show that in the limit of infinitely slow acceleration of the wave, these highly resonant particles disappear and the response

  18. Evaluation of performance and characteristics of long-pulse Hall-ion thrusters utilizing a power source based on CDL capacitor technology

    NASA Astrophysics Data System (ADS)

    Hrbud, Ivana

    1997-09-01

    In 1994, NASA initiated the New Millennium Program to identify, promote, and to fund research and development of specific advanced technologies for space and Earth exploration in the 21st century. NASA proposes the frequent launch of affordable missions to be accomplished by small, low-cost spacecraft which will employ the advantages of high-power, high-efficiency thrusters without producing excessive demand on the spacecraft's power train. Since the power systems of these spacecraft will be power limited, more efficient propulsion may be achieved by pulsed mode operation of a Hall-ion thruster. The goals of this research are (1) to prove the concept feasibility of a direct-drive electric propulsion system, and (2) to evaluate the performance and characteristics of a Russian TAL (Thruster with Anode Layer) operating in a long-pulse mode, powered by a capacitor-based power source developed at Space Power Institute. The TAL, designated D-55, is characterized by an external acceleration zone and is powered by a unique chemical double layer (CDL) capacitor bank with a capacitance of 4 F at a charge voltage of 400 V. Performance testing of this power supply on the TAL was conducted at NASA Lewis Research Center in Cleveland, OH. Direct thrust measurements of the TAL were obtained at CDL power levels ranging from 450 to 1750 W. The specific impulse encompassed a range from 1150 s to 2200 s, yielding thruster system efficiencies between 50 and 60%. Preliminary mission analysis of the CDL direct-drive concept and other electric propulsion options was performed for the ORACLE spacecraft in 6am/6pm and 12am/12pm, 300 km sun-synchronous orbits. The direct-drive option was competitive with the other systems by increasing available net mass between 5 and 42% and reducing two-year system wet mass between 18 and 63%. Overall, the electric propulsion power requirements for the satellite were reduced between 57 and 91% depending on the orbit evaluated. The direct-drive, CDL

  19. 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.

  20. Temperature and current accelerated lifetime conditions and testing of laser diodes for ESA BepiColombo space mission

    NASA Astrophysics Data System (ADS)

    Klumel, Genady; Karni, Yoram; Cohen, Shalom; Rech, Markus; Weidlich, Kai

    2011-03-01

    System designers and end users of diode pumped solid state lasers often require knowledge of the operability limits of QCW laser diode pump sources and their predicted reliability performance as a function of operating conditions. Accelerated ageing at elevated temperatures, duty cycles and/or currents allows extended lifetime testing of diode stacks to be executed on compressed timescales with high confidence. We present a novel, time-efficient technique for the determination of accelerated lifetime test conditions using degradation rate data, rather than the traditionally used failures against time data. To assess the effect of thermally accelerated ageing, 4 groups of 4 stacks each were operated for 60 million pulses at different temperature stress levels by varying the pulse repetition rate from 100Hz to 250Hz. The measured power degradation rates fitted to an Arrhenius type model, result in activation energy of 0.47- 0.74eV, apparently indicating two thermally activated degradation modes with different activation energies. Similarly, for current accelerated ageing, another 4 groups of 4 stacks were tested at operation currents from 120A to 150A. The optical power degradation rates due to current stress follow a power law behavior with a current acceleration factor of 1.7. The obtained acceleration parameters allowed considerable reduction of the lifetime test duration, which would have otherwise taken an unacceptably long time under nominal operating conditions. The successful results of the accelerated lifetime have been a major milestone enabling qualification of SCD stacks as pump sources for the laser altimeter in ESA Bepi-Colombo space mission. The presented reliability analysis allows life test qualification programs to be accelerated for generic QCW stacks and their lifetime to be predicted in various operating environments.

  1. Control of the Electric Field Profile in the Hall Thruster

    SciTech Connect

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

    2000-10-05

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

  2. 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

    Electric thrusters are being considered for a variety of space missions because of the significant propellant savings that result from the use of high performance, electric propulsion technologies, Propellant mass savings reduces spacecraft launch requirements and increases mission lifetime and payload. The impact of electric thruster plasma plumes on microwave signal propagation however is an important spacecraft integration concern. Arcjets were the first electric thrusters to be considered for operational missions. Ling, et al., studied the effect of arcjet plumes on propagation. Arcjets produce a lightly ionized plume and Ling's analysis predicted that the plume would have a negligible effect on communication. Plumes from the higher performance ion thrusters being developed exhibit higher ionization levels, plasma temperatures and particle velo@ities than arcjets. Therefore, there was a need to assess the impact due to these plumes. To address this need, the authors designed and performed a series of experiments to examine propagation effects of plumes. The challenge with these experiments was that they had to be performed in the operational environment of the thruster. Therefore, the experiments were conducted inside a metal chamber which could be depressurized to simulate a near vacuum condition of space. The metal chamber presents a potential large source of error to the propagation measurements due to the corruption of the desired data by multiple wall reflections within the chamber. This chamber effect was minimized by employing a pulsed-continuous wave transmitter and receiver system. This system, based on an HP8510 Network Analyzer, uses external hardware time gating to eliminate the clutter of the spurious reflections. Additionally, high gain antennas were used in the measurements to ensure that minimal amounts of energy ",ere transmitted/received in undesirable directions. The measurements took place in Vacuum Facility 5 of the Electric Propulsion

  3. 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

    Electric thrusters are being considered for a variety of space missions because of the significant propellant savings that result from the use of high performance, electric propulsion technologies. Propellant mass savings reduces spacecraft launch requirements and increases mission lifetime and payload. The impact of electric thruster plasma plumes on microwave signal propagation however is an important spacecraft integration concern. Arcjets were the first electric thrusters to be considered for operational missions. Ling, et al. studied the effect of arcjet plumes on propagation. Arcjets produce a lightly ionized plume and Ling's analysis predicted that the plume would have a negligible effect on communication. Plumes from the higher performance ion thrusters being developed exhibit higher ionization levels, plasma temperatures and particle velocities than arcjets. Therefore, there was a need to assess the impact due to these plumes. To address this need, the authors designed and performed a series of experiments to examine propagation effects of plumes. The challenge with these experiments was that they had to be performed in the operational environment of the thruster. Therefore, the experiments were conducted inside a metal chamber which could be depressurized to simulate a near vacuum condition of space. The metal chamber presents a potential large source of error to the propagation measurements due to the corruption of the desired data by multiple wall reflections within the chamber. This chamber effect was minimized by employing a pulsed-continuous wave transmitter and receiver system. This system based on an HP8510 Network Analyzer, uses external hardware time gating to eliminate the clutter of the spurious reflections. Additionally, high gain antennas were used in the measurements to ensure that minimal amounts of energy were transmitted/received in undesirable directions. The measurements took place in Vacuum Facility 5 of the Electric Propulsion

  4. Coaxial plasma thrusters for high specific impulse propulsion

    NASA Technical Reports Server (NTRS)

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

    1991-01-01

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

  5. Investigations of Probe Induced Perturbations in a Hall Thruster

    SciTech Connect

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

    2002-08-12

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

  6. Factors Affecting the Efficiency of Krypton Hall Thrusters

    NASA Technical Reports Server (NTRS)

    Hofer, Richard R.; Peterson, Peter Y.; Jacobson, David T.; Manzella, David M.

    2004-01-01

    The krypton-fueled Hall thruster offers the possibility of high-specific impulse and long lifetime. NASA's series of Hall thrusters have demonstrated krypton efficiencies only 5 - 15% less than xenon. Larger thrusters have smaller differences in efficiency. Plasma measurements have demonstrated that efficiency is reduced due to a decrease in mass utilization. Current efforts are considering the implications of these results, and how design changes can be made to increase the efficiency of krypton Hall thrusters.

  7. A nuclear driven metallic vapor MHD coupled with MPD thrusters

    NASA Technical Reports Server (NTRS)

    Anghaie, Samim; Kumar, Ratan

    1991-01-01

    Nuclear energy as a source of power for space missions, represents an enabling technology for advanced and ambitious space applications. Nuclear fuel in a gaseous or liquid form has been configured as a promising and practical candidate in this regard. The present study investigates and performs a feasibility analysis of an innovative concept for space power generation and propulsion. The system embodies a conceptual nuclear reactor with an MHD generator and coupled to MPD thrusters. The reactor utilizes liquid uranium in droplet form as fuel and superheated metallic vapor as the working fluid. This ultrahigh temperature vapor core reactor brings forward varied and challenging technical issues, and it has been addressed to in this paper. A parametric study of the conceived system has been performed in a qualitative and quantitative manner. Preliminary results show enough promise for further indepth analysis of this novel system.

  8. Electrostatic acceleration of helicon plasma using a cusped magnetic field

    SciTech Connect

    Harada, S.; Baba, T.; Uchigashima, A.; Iwakawa, A.; Sasoh, A.; Yokota, S.; Yamazaki, T.; Shimizu, H.

    2014-11-10

    The electrostatic acceleration of helicon plasma is investigated using an electrostatic potential exerted between the ring anode at the helicon source exit and an off-axis hollow cathode in the downstream region. In the downstream region, the magnetic field for the helicon source, which is generated by a solenoid coil, is modified using permanent magnets and a yoke, forming an almost magnetic field-free region surrounded by an annular cusp field. Using a retarding potential analyzer, two primary ion energy peaks, where the lower peak corresponds to the space potential and the higher one to the ion beam, are detected in the field-free region. Using argon as the working gas with a helicon power of 1.5 kW and a mass flow rate of 0.21 mg/s, the ion beam energy is on the order of the applied acceleration voltage. In particular, with an acceleration voltage lower than 150 V, the ion beam energy even exceeds the applied acceleration voltage by an amount on the order of the electron thermal energy at the exit of the helicon plasma source. The ion beam energy profile strongly depends on the helicon power and the applied acceleration voltage. Since by this method the whole working gas from the helicon plasma source can, in principle, be accelerated, this device can be applied as a noble electrostatic thruster for space propulsion.

  9. Improvements in Modeling Thruster Plume Erosion Damage to Spacecraft Surfaces

    NASA Technical Reports Server (NTRS)

    Soares, Carlos; Olsen, Randy; Steagall, Courtney; Huang, Alvin; Mikatarian, Ron; Myers, Brandon; Koontz, Steven; Worthy, Erica

    2015-01-01

    Spacecraft bipropellant thrusters impact spacecraft surfaces with high speed droplets of unburned and partially burned propellant. These impacts can produce erosion damage to optically sensitive hardware and systems (e.g., windows, camera lenses, solar cells and protective coatings). On the International Space Station (ISS), operational constraints are levied on the position and orientation of the solar arrays to mitigate erosion effects during thruster operations. In 2007, the ISS Program requested evaluation of erosion constraint relief to alleviate operational impacts due to an impaired Solar Alpha Rotary Joint (SARJ). Boeing Space Environments initiated an activity to identify and remove sources of conservatism in the plume induced erosion model to support an expanded range of acceptable solar array positions ? The original plume erosion model over-predicted plume erosion and was adjusted to better correlate with flight experiment results. This paper discusses findings from flight experiments and the methodology employed in modifying the original plume erosion model for better correlation of predictions with flight experiment data. The updated model has been successful employed in reducing conservatism and allowing for enhanced flexibility in ISS solar array operations.

  10. Dual-throat thruster thermal model. Final report

    SciTech Connect

    Ewen, R.L.; Obrien, C.J.; Matthews, L.W.

    1986-08-01

    The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

  11. SERT 2 1979 extended flight thruster system performance

    NASA Technical Reports Server (NTRS)

    Kerslake, W. R.; Ignaczak, L. R.

    1979-01-01

    Steady state tests of the thruster 2 system on the SERT 2 spacecraft are presented. A direct thrust measurement was obtained for the ion thruster during operations to increase the spacecraft spin rate to maintain spacecraft attitude stability. The continued restart tests of thruster 1 and a report on the general status of all spacecraft systems including the main solar array are presented.

  12. Parametric Investigations of Non-Conventional Hall Thruster

    SciTech Connect

    Raitses, Y.; Fisch, N.J.

    2001-01-12

    Hall thrusters might better scale to low power with non-conventional geometry. A 9 cm cylindrical, ceramic-channel, Hall thruster with a cusp-type magnetic field distribution has been investigated. It exhibits discharge characteristics similar to conventional coaxial Hall thrusters, but does not expose as much channel surface. Significantly, its operation is not accompanied by large amplitude discharge low frequency oscillations.

  13. Optical Characterization of Component Wear and Near-Field Plasma of the Hermes Thruster

    NASA Technical Reports Server (NTRS)

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

    2015-01-01

    Optical emission spectral (OES) data are presented which correlate trends in sputtered species and the near-field plasma with the Hall-Effect Rocket with Magnetic Shielding (HERMeS) thruster operating condition. The relative density of singly-ionized xenon (Xe II) is estimated using a collisional-radiative model. OES data were collected at three radial and several axial locations downstream of the thruster's exit plane. These data were deconvolved to show the structure for the near-field plasma as a function of thruster operating condition. The magnetic field is shown to have a much greater affect on plasma structure than the discharge voltage with the primary ionization/acceleration zone boundary being similar for all nominal operating voltages at constant power. OES measurement of sputtered boron shows that the HERMeS thruster is magnetically shielded across its operating envelope. Preliminary assessment of carbon sputtered from the keeper face suggest it increases significantly with operating voltage, but the uncertainty associated with these measurements is very high.

  14. Kinetic particle simulation of discharge and wall erosion of a Hall thruster

    SciTech Connect

    Cho, Shinatora; Komurasaki, Kimiya; Arakawa, Yoshihiro

    2013-06-15

    The primary lifetime limiting factor of Hall thrusters is the wall erosion caused by the ion induced sputtering, which is predominated by dielectric wall sheath and pre-sheath. However, so far only fluid or hybrid simulation models were applied to wall erosion and lifetime studies in which this non-quasi-neutral and non-equilibrium area cannot be treated directly. Thus, in this study, a 2D fully kinetic particle-in-cell model was presented for Hall thruster discharge and lifetime simulation. Because the fully kinetic lifetime simulation was yet to be achieved so far due to the high computational cost, the semi-implicit field solver and the technique of mass ratio manipulation was employed to accelerate the computation. However, other artificial manipulations like permittivity or geometry scaling were not used in order to avoid unrecoverable change of physics. Additionally, a new physics recovering model for the mass ratio was presented for better preservation of electron mobility at the weakly magnetically confined plasma region. The validity of the presented model was examined by various parametric studies, and the thrust performance and wall erosion rate of a laboratory model magnetic layer type Hall thruster was modeled for different operation conditions. The simulation results successfully reproduced the measurement results with typically less than 10% discrepancy without tuning any numerical parameters. It is also shown that the computational cost was reduced to the level that the Hall thruster fully kinetic lifetime simulation is feasible.

  15. Perturbations induced by electrostatic probe in the discharge of Hall thrusters.

    PubMed

    Grimaud, L; Pétin, A; Vaudolon, J; Mazouffre, S

    2016-04-01

    Emissive and Langmuir probes are two widely used plasma diagnostic techniques that, when used properly, give access to a wide range of information on the plasma's ions and electrons. We show here that their use in small and medium power Hall thrusters produces large perturbations in the discharge characteristics. Potential measurements performed by both probes and non-invasive Laser Induced Fluorescence (LIF) spectroscopy highlight significant discrepancies in the discharge profile. This phenomenon is observed both in the 200 W and the 1.5 kW-class thrusters. In order to have a better understanding of these perturbations, ion velocity distribution functions are acquired by LIF spectroscopy at different positions in the smaller thruster, with and without the probes. Emissive probes are shown to produce the biggest perturbation, shifting the acceleration region upstream. The probe insertion is also shown to have significant effect on both the average discharge current, increasing it by as much as 30%, and its harmonic content in both amplitude and spectrum. These perturbations appear as the probe tip passes a threshold located between 0 and 5 mm downstream of the thruster exit plane.

  16. Perturbations induced by electrostatic probe in the discharge of Hall thrusters

    NASA Astrophysics Data System (ADS)

    Grimaud, L.; Pétin, A.; Vaudolon, J.; Mazouffre, S.

    2016-04-01

    Emissive and Langmuir probes are two widely used plasma diagnostic techniques that, when used properly, give access to a wide range of information on the plasma's ions and electrons. We show here that their use in small and medium power Hall thrusters produces large perturbations in the discharge characteristics. Potential measurements performed by both probes and non-invasive Laser Induced Fluorescence (LIF) spectroscopy highlight significant discrepancies in the discharge profile. This phenomenon is observed both in the 200 W and the 1.5 kW-class thrusters. In order to have a better understanding of these perturbations, ion velocity distribution functions are acquired by LIF spectroscopy at different positions in the smaller thruster, with and without the probes. Emissive probes are shown to produce the biggest perturbation, shifting the acceleration region upstream. The probe insertion is also shown to have significant effect on both the average discharge current, increasing it by as much as 30%, and its harmonic content in both amplitude and spectrum. These perturbations appear as the probe tip passes a threshold located between 0 and 5 mm downstream of the thruster exit plane.

  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. Investigation of a pulsed electrothermal thruster system

    NASA Technical Reports Server (NTRS)

    Burton, R. L.; Goldstein, S. A.; Hilko, B. K.; Tidman, D. A.; Winsor, N. K.

    1984-01-01

    The performance of an ablative wall Pulsed Electrothermal (PET) thruster is accurately characterized on a calibrated thrust stand, using polyethylene propellant. The thruster is tested for four configurations of capillary length and pulse length. The exhaust velocity is determined with twin time-of-flight photodiode stagnation probes, and the ablated mass is measured from the loss over ten shots. Based on the measured thrust impulse and the ablated mass, the specific impulse varies from 1000 to 1750 seconds. The thrust to power varies from .05 N/kW (quasi-steady mode) to .10 N/kW (unsteady mode). The thruster efficiency varies from .56 at 1000 seconds to .42 at 1750 seconds. A conceptual design is presented for a 40 kW PET propulsion system. The point design system performance is .62 system efficiency at 1000 seconds specific impulse. The system's reliability is enhanced by incorporating 20, 20 kW thruster modules which are fired in pairs. The thruster design is non-ablative, and uses water propellant, from a central storage tank, injected through the cathode.

  19. Physics of Phase Space Matching for Staging Plasma and Traditional Accelerator Components Using Longitudinally Tailored Plasma Profiles

    NASA Astrophysics Data System (ADS)

    Xu, X. L.; Hua, J. F.; Wu, Y. P.; Zhang, C. J.; Li, F.; Wan, Y.; Pai, C.-H.; Lu, W.; An, W.; Yu, P.; Hogan, M. J.; Joshi, C.; Mori, W. B.

    2016-03-01

    Phase space matching between two plasma-based accelerator (PBA) stages and between a PBA and a traditional accelerator component is a critical issue for emittance preservation. The drastic differences of the transverse focusing strengths as the beam propagates between stages and components may lead to a catastrophic emittance growth even when there is a small energy spread. We propose using the linear focusing forces from nonlinear wakes in longitudinally tailored plasma density profiles to control phase space matching between sections with negligible emittance growth. Several profiles are considered and theoretical analysis and particle-in-cell simulations show how these structures may work in four different scenarios. Good agreement between theory and simulation is obtained, and it is found that the adiabatic approximation misses important physics even for long profiles.

  20. Physics of Phase Space Matching for Staging Plasma and Traditional Accelerator Components Using Longitudinally Tailored Plasma Profiles.

    PubMed

    Xu, X L; Hua, J F; Wu, Y P; Zhang, C J; Li, F; Wan, Y; Pai, C-H; Lu, W; An, W; Yu, P; Hogan, M J; Joshi, C; Mori, W B

    2016-03-25

    Phase space matching between two plasma-based accelerator (PBA) stages and between a PBA and a traditional accelerator component is a critical issue for emittance preservation. The drastic differences of the transverse focusing strengths as the beam propagates between stages and components may lead to a catastrophic emittance growth even when there is a small energy spread. We propose using the linear focusing forces from nonlinear wakes in longitudinally tailored plasma density profiles to control phase space matching between sections with negligible emittance growth. Several profiles are considered and theoretical analysis and particle-in-cell simulations show how these structures may work in four different scenarios. Good agreement between theory and simulation is obtained, and it is found that the adiabatic approximation misses important physics even for long profiles.