WorldWideScience

Sample records for repetitively-pulsed electric thrusters

  1. Non-contact thrust stand calibration method for repetitively pulsed electric thrusters.

    Science.gov (United States)

    Wong, Andrea R; Toftul, Alexandra; Polzin, Kurt A; Pearson, J Boise

    2012-02-01

    A thrust stand calibration technique for use in testing repetitively pulsed electric thrusters for in-space propulsion has been developed and tested using a modified hanging pendulum thrust stand. In the implementation of this technique, current pulses are applied to a solenoid to produce a pulsed magnetic field that acts against a permanent magnet mounted to the thrust stand pendulum arm. The force on the magnet is applied in this non-contact manner, with the entire pulsed force transferred to the pendulum arm through a piezoelectric force transducer to provide a time-accurate force measurement. Modeling of the pendulum arm dynamics reveals that after an initial transient in thrust stand motion the quasi-steady average deflection of the thrust stand arm away from the unforced or "zero" position can be related to the average applied force through a simple linear Hooke's law relationship. Modeling demonstrates that this technique is universally applicable except when the pulsing period is increased to the point where it approaches the period of natural thrust stand motion. Calibration data were obtained using a modified hanging pendulum thrust stand previously used for steady-state thrust measurements. Data were obtained for varying impulse bit at constant pulse frequency and for varying pulse frequency. The two data sets exhibit excellent quantitative agreement with each other. The overall error on the linear regression fit used to determine the calibration coefficient was roughly 1%.

  2. Non-Contact Thrust Stand Calibration Method for Repetitively-Pulsed Electric Thrusters

    Science.gov (United States)

    Wong, Andrea R.; Toftul, Alexandra; Polzin, Kurt A.; Pearson, J. Boise

    2011-01-01

    A thrust stand calibration technique for use in testing repetitively-pulsed electric thrusters for in-space propulsion has been developed and tested using a modified hanging pendulum thrust stand. In the implementation of this technique, current pulses are applied to a solenoidal coil to produce a pulsed magnetic field that acts against the magnetic field produced by a permanent magnet mounted to the thrust stand pendulum arm. The force on the magnet is applied in this non-contact manner, with the entire pulsed force transferred to the pendulum arm through a piezoelectric force transducer to provide a time-accurate force measurement. Modeling of the pendulum arm dynamics reveals that after an initial transient in thrust stand motion the quasisteady average deflection of the thrust stand arm away from the unforced or zero position can be related to the average applied force through a simple linear Hooke s law relationship. Modeling demonstrates that this technique is universally applicable except when the pulsing period is increased to the point where it approaches the period of natural thrust stand motion. Calibration data were obtained using a modified hanging pendulum thrust stand previously used for steady-state thrust measurements. Data were obtained for varying impulse bit at constant pulse frequency and for varying pulse frequency. The two data sets exhibit excellent quantitative agreement with each other as the constant relating average deflection and average thrust match within the errors on the linear regression curve fit of the data. Quantitatively, the error on the calibration coefficient is roughly 1% of the coefficient value.

  3. Repetitively Pulsed Electric Laser Acoustic Studies. Volume 1.

    Science.gov (United States)

    1983-09-01

    INGARD ET AL. SEP 83 UNCLASSIFIED APHAL-IR-83-2858-VOL-1 F336i5 86-C 2848 F/ 0/ 8, EEEmohEEEomhiE EohEEmhohEEEEE mhhhmmomhhlm...TR-83-2058, Vol 9, 0 REPETITIVELY PULSED ELECTRIC LASER ACOUSTIC STUDIES Volume I K. U. INGARD , CHARLES F. MCMILLAN uDEPARTMENT OF AERONAUTICS AND...CONTRACT OR GRANT NUMBER(s) K.U. Ingard and Charles F. McMillan F33615.80-C-2040 9. PERFORMING ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT. PROJECT

  4. NASA HERMeS Hall Thruster Electrical Configuration Characterization

    Science.gov (United States)

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

    2015-01-01

    The NASA Hall Effect Rocket with Magnetic Shielding (HERMeS) 12.5 kW Technology Demonstration Unit-1 (TDU-1) Hall thruster has been the subject of extensive technology maturation in preparation for development into a flight ready propulsion system. Part of the technology maturation was to test the TDU-1 thruster in several ground based electrical configurations to assess the thruster robustness and suitability to successful in-space operation. The ground based electrical configuration testing has recently been demonstrated as an important step in understanding and assessing how a Hall thruster may operate differently in-space compared to ground based testing, and to determine the best configuration to conduct development and qualification testing. This paper describes the electrical configuration testing of the HERMeS TDU-1 Hall thruster in NASA Glenn Research Center's Vacuum Facility 5. The three electrical configurations examined were 1) thruster body tied to facility ground, 2) thruster floating, and 3) thruster body electrically tied to cathode common. The HERMeS TDU-1 Hall thruster was also configured with two different exit plane boundary conditions, dielectric and conducting, to examine the influence on the electrical configuration characterization.

  5. 电激励重复频率非链式HF激光器%Electrically initiated repetitive-pulsed non-chain HF lasers

    Institute of Scientific and Technical Information of China (English)

    易爱平; 刘晶儒; 唐影; 黄珂; 黄欣; 于力; 马连英

    2011-01-01

    采用电子束和气体放电两种激励方式开展重复频率非链式HF激光研究.基于全固态脉冲功率源SPG200建立了重复频率HF实验装置,探索了产生重频大面积均匀电子束的技术途径,利用法拉第筒对进入激光气室的电子束的轴向均匀性进行了诊断,开展了激光器输出特性研究和重频实验调试,在C2 H6:SF6=0.035,总气压为35 kPa时,激光器输出能量最大约为4.8 mJ,并实现了最高30 Hz的HF激光稳定输出.采用峰化电容及紫外自动预电离技术设计研制了放电激励重频HF激光器,研究了SF6气体放电特性和重频运行稳定性.研究发现SF6气体放电具有典型的辉光放电、电压维持和电弧放电三阶段特征.在充电电压为28 kV,总气压为12 kPa,C2H6含量为8%时,放电激励HF激光器最大脉冲能量约 600 mJ,比能量输出达到8.5J/1,激光器的电光转换效率约为2.5%.该激光器在1-50 Hz实现了重频输出,首脉冲能量>500 mJ,在10 Hz时稳定输出能量约为200 mJ,%The repetitive-pulsed non-chain HF lasers initiated by e-beams and fast discharges were studied respectively. The e-beam initiated HF laser characterized by an all-solid-state generator was developed. A large area repetitive uniform e-beam was obtained and the Faraday cup was used to diagnose the uniformity of e-beam into the laser gas cell. The output characteristics of the laser were studied and the maximal energy of 4.8 mJ was obtained with gas mixture C2 H6: SF6 =0.035. The e-beam initiated HF laser could operate at the repetition rate of 30 Hz. The discharge initiated non-chain HF laser was designed by using the technologies of peaking capacitor and UV pre-ionizing. The discharge characteristics and stability of SF6 gas in repetition mode were studied. The results show that the discharge includes main discharge (volume discharge) and arc discharge, between which the voltage maintains and no evident discharge is visible. The maximal output

  6. NASA HERMeS Hall Thruster Electrical Configuration Characterization

    Science.gov (United States)

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

    2016-01-01

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

  7. Gas-dynamic perturbations in an electric-discharge repetitively pulsed DF laser and the role of He in their suppression

    Energy Technology Data Exchange (ETDEWEB)

    Evdokimov, P A; Sokolov, D V [Russian Federal Nuclear Center ' All-Russian Research Institute of Experimental Physics' , Sarov, Nizhnii Novgorod region (Russian Federation)

    2015-11-30

    The gas-dynamic perturbations in a repetitively pulsed DF laser are studied using a Michelson interferometer. Based on the analysis of experimental data obtained in two experimental sets (working medium without buffer gas and with up to 90% of He), it is concluded that such phenomena as isentropic expansion of a thermal plug, gas heating by shock waves and resonance acoustic waves do not considerably decrease the upper limit of the pulse repetition rate below a value determined by the time of the thermal plug flush out of the discharge gap. It is suggested that this decrease for a DF laser with the SF{sub 6} – D{sub 2} working mixture is caused by the development of overheat instability due to an increased energy deposition into the near-electrode regions and to the formation of electrode shock waves. Addition of He to the active media of the DF laser changes the discharge structure and improves its homogeneity over the discharge gape cross section, thus eliminating the reason for the development of this instability. A signification dilution of the active medium of a DF laser with helium up to the atmospheric pressure allowed us to achieve the limiting discharge initiation frequencies with the active medium replacement ratio K ∼ 1. (active media)

  8. Carbon Nanotube Based Electric Propulsion Thruster with Low Power Consumption Project

    Data.gov (United States)

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

  9. Spectroscopic Investigation of a Repetitively-Pulsed Nanosecond Discharge

    Science.gov (United States)

    Yee, Benjamin T.

    This work reports on an investigation of a repetitively-pulsed nanosecond discharge (RPND) in helium over a range of 0.3-16.0 Torr. The discharge was studied experimentally via laser-absorption spectroscopy and opticals emission spectroscopy measurements. In concert with the experimental campaign, a global model of a helium plasma was developed with the aid of particle-in-cell simulations. The global model was then used to predict the population kinetics and emissions of the RPND. Synthesis of the results provided new data and insights on the development of the RPND. Among the results were direct measurements of the triplet metastable states during the excitation period. This period was found to be unexpectedly long at low pressures (less than or equal to 1.0 Torr), suggesting an excess in high-energy electrons as compared to an equilibrium distribution. Other phenomena such as a prominent return stroke and additional energy deposition by reflections in the transmission line were also identified. Estimates of the electric field and electron temperatures were obtained for several conditions. Furthermore, several optical methods for electron temperature measurement were evaluated for application to the discharge. Based on the global model simulations, the coronal model was found to apply to the line ratio of the 33S-23Po and 31S-2 1Po transitions, however further work is needed to ascertain its applicability to experimental discharges. These results provide new insight on the development of the repetitively-pulsed nanosecond discharge. Specifically, they reveal new information about the excited state dynamics within the discharge, the non-equilibrium nature of its electrons, and several avenues for future studies. This study extends the present understanding of repetitively-pulsed discharges, and advances the knowledge of energy coupling between electric fields and plasmas.

  10. Carbon Nanotube Based Electric Propulsion Thruster with Low Power Consumption Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Field emission electric propulsion (FEEP) thrusters have gained considerable attention for spacecrafts disturbance compensation because of excellent characteristics....

  11. Simple filtered repetitively pulsed vacuum arc plasma source

    Science.gov (United States)

    Chekh, Yu.; Zhirkov, I. S.; Delplancke-Ogletree, M. P.

    2010-02-01

    A very simple design of cathodic filtered vacuum arc plasma source is proposed. The source without filter has only four components and none of them require precise machining. The source operates in a repetitively pulsed regime, and for laboratory experiments it can be used without water cooling. Despite the simple construction, the source provides high ion current at the filter outlet reaching 2.5% of 400 A arc current, revealing stable operation in a wide pressure range from high vacuum to oxygen pressure up to more than 10-2 mbar. There is no need in complicated power supply system for this plasma source, only one power supply can be used to ignite the arc, to provide the current for the arc itself, to generate the magnetic field in the filter, and provide its positive electric biasing without any additional high power resistance.

  12. Optimal electric potential profile in a collisional magnetized thruster

    Science.gov (United States)

    Fruchtman, Amnon; Makrinich, Gennady

    2016-10-01

    A major figure of merit in propulsion in general and in electric propulsion in particular is the thrust per unit of deposited power, the ratio of thrust over power. We have recently demonstrated experimentally and theoretically that for a fixed deposited power in the ions, the momentum delivered by the electric force is larger if the accelerated ions collide with neutrals during the acceleration. As expected, the higher thrust for given power is achieved for a collisional plasma at the expense of a lower thrust per unit mass flow rate. Operation in the collisional regime can be advantageous for certain space missions. We analyze a Hall thruster configuration in which the flow is only weakly ionized but there are frequent ion-neutral collisions. With a variational method we seek an electric potential profile that maximizes thrust over power. We then examine what radial magnetic field profile should determine such a potential profile. Supported by the Israel Science Foundation Grant 765/11.

  13. Electric Propulsion Cables For Milli-Newton Thrusters

    Science.gov (United States)

    Jakob, Manfred; Bertrand, Arnaud; El-Idrissi, Mohamed; Schaper, Wolfgang, , Dr.

    2011-10-01

    AXON' Kabel GmbH, is developing and manufacturing cables and connectors up to complete interconnect systems for all types of applications needed in Space. As a request from ESA, AXON has developed a new generation of cables suitable for current and future applications to feed electric propulsion thruster systems in spacecraft with electric power. Under this project the main objectives were to find and select materials for the composition to produce a cable withstanding quite strongrequirements for operating temperature, radiation resistance, high voltage application and in variants to various current ratings (A); the cable construction will also include ESD immunisation. The paper will summarise the specification achieved and will give an overview on the test results with the prototype cables.

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

    Science.gov (United States)

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

    2016-10-01

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

  15. Repetitively Pulsed Backward-Wave Oscillator Investigations

    Science.gov (United States)

    1994-03-31

    and the FE phase by applying a’ Y pulsed electric field , Fig. 1. Sawyer-Tower circuit for displaying (4) partial reversal of P. inside the P-E...at temper- a pulsed electric field to switch the material atures up to the Curie temperature. Tests on into the PE or APE phase. With this combina- a

  16. Improvement of Polytetrafluoroethylene Surface Energy by Repetitive Pulse Non-Thermal Plasma Treatment in Atmospheric Air%Improvement of Polytetrafluoroethylene Surface Energy by Repetitive Pulse Non-Thermal Plasma Treatment in Atmospheric Air

    Institute of Scientific and Technical Information of China (English)

    杨国清; 张冠军; 张文元

    2011-01-01

    Improvement of polytetrafluoroethylene surface energy by non-thermal plasma treatment is presented, using a nanosecond-positive-edge repetitive pulsed dielectric barrier discharge generator in atmospheric air. The electrical parameters including discharging power, peak and density of micro-discharge current were calculated, and the electron energy was estimated. Surface treatment experiments of polytetrafluoroethylene films were conducted for both different applied voltages and different treating durations. Results show that the surface energy of polytetrafluoroethylene film could be improved to 40 mJ/m2 or more by plasma treatment. Surface roughness measurement and surface X-ray photoelectron spectroscopy analysis indicate that there are chemical etching and implantation of polar oxygen groups in the sample surface treating process, resulting in the improvement of the sample surface energy. Compared with an AC source of 50 Hz, the dielectric barrier discharges generated by a repetitive pulsed source could provide higher peak power, lower mean power, larger micro-discharge current density and higher electron energy. Therefore, with the same applied peak voltage and treating duration, the improvement of polytetrafluoroethylene surface energy using repetitive pulsed plasma is more effective, and the plasma treatment process based on repetitive pulsed dielectric barrier discharges in air is thus feasible and applicable.

  17. Magnetoplasmadynamic electric propulsion thruster behavior at the hundred megawatt level

    Science.gov (United States)

    Marriott, Darin William

    Characteristic measurements were made of a hundred megawatt modified helium inverse pinch switch and compared against numerical modeling and theoretically expected behavior. Thruster voltage was measured for currents between three and three hundred kilo amps and for mass flow rates between 0.96 and 40 grams per second. From that, characteristic voltage, power, and resistance curves were generated. Electron temperature measurements made inside the plasma flow using triple Langmuir probes were found to be between three and thirty electron volts. General expected MPD thruster behavior, such as decreasing resistance with increasing mass flow rate, were confirmed. The quasi steady assumption was studied between 1.5 and 1.7 milliseconds and found to be appropriate. A theoretical model, based on integrating the magnetic field to determine thrust, as for an MPD thruster, was used to estimate fall voltages, pumping coefficients, and specific impulse. An empirical model for thruster voltage was then created to estimate the behavior of voltage as a function of the similarity parameter. The two models were then put together and found to be self consistent with the experimental data. Three sources of power loss were estimated given the experimental and theoretical model. The power lost due to fall voltage mechanisms was calculated from the theoretical model and the input current as a function of time. The ionization losses were estimated using a worst case scenario of complete double ionization of the input helium mass flow rate as a function of time. Thermal losses were calculated from the electron temperature and the input mass flow rate. Total temperature, specific impulse, and efficiency measurements were all presented as a function of a similarity parameter in line with MPD theory. Basic MPD thruster behavior was confirmed. Suggestions were made for future continuation of the project.

  18. Repetitively pulsed Fe: ZnSe laser with an average output power of 20 W at room temperature of the polycrystalline active element

    Science.gov (United States)

    Velikanov, S. D.; Gavrishchuk, E. M.; Zaretsky, N. A.; Zakhryapa, A. V.; Ikonnikov, V. B.; Kazantsev, S. Yu.; Kononov, I. G.; Maneshkin, A. A.; Mashkovskii, D. A.; Saltykov, E. V.; Firsov, K. N.; Chuvatkin, R. S.; Yutkin, I. M.

    2017-05-01

    The energy and spectral-temporal characteristics of a Fe : ZnSe laser operating in pulsed and repetitively pulsed regimes are studied at room temperature of the polycrystalline active element. The crystal was pumped by a nonchain electric-discharge HF laser. The energy of the Fe : ZnSe laser in a single-pulse regime was 1.67 J at the slope efficiency with respect to the absorbed and incident energy of ∼43% and ∼27%, respectively. In a repetitively pulsed regime with a pulse repetition rate of 20 Hz and an efficiency with respect to the absorbed power of ∼40%, the average laser power was ∼20 W with an individual pulse energy of ∼1 J. The possibility of increasing the average power of the repetitively pulsed Fe : ZnSe laser at room temperature is discussed.

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

    Science.gov (United States)

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

    1995-01-01

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

  20. Development of a compact and reliable repetitively pulsed Xe Cl (308 nm) excimer laser

    Indian Academy of Sciences (India)

    N S Benerji; N Varshnay; J K Mittal

    2013-02-01

    Development and operation characteristics of a repetitively pulsed UV spark pre-ionized XeCl(Xenon Chloride) excimer laser is described. The laser uses discharge pumped C–C charge transfer excitation. A compact gas circulation loop was adopted to achieve high repetition rate operation. The laser generates optical pulses of energy 150 mJ at 150 Hz reliably. The electrical to optical conversion efficiency obtained is 1%. The laser pulse duration is ∼8 nS (FWHM). The single fill gas lifetime have been found to be 2 × 106 shots for 20% reduction of energy without any halogen injection. The system is compact and reliable.

  1. Atmospheric Breathing Electric Thruster for Planetary Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — This study will investigate the development of an atmosphere-breathing electric propulsion solar-powered vehicle to explore planets such as Mars. The vehicle would...

  2. Power neodymium-glass amplifier of a repetitively pulsed laser

    Energy Technology Data Exchange (ETDEWEB)

    Vinogradov, Aleksandr V; Gaganov, V E; Garanin, Sergey G; Zhidkov, N V; Krotov, V A; Martynenko, S P; Pozdnyakov, E V; Solomatin, I I [Russian Federal Nuclear Center ' All-Russian Research Institute of Experimental Physics' , Sarov, Nizhnii Novgorod region (Russian Federation)

    2011-11-30

    A neodymium-glass diode-pumped amplifier with a zigzag laser beam propagation through the active medium was elaborated; the amplifier is intended for operation in a repetitively pulsed laser. An amplifier unit with an aperture of 20 Multiplication-Sign 25 mm and a {approx}40-cm long active medium was put to a test. The energy of pump radiation amounts to 140 J at a wavelength of 806 nm for a pump duration of 550 {mu}s. The energy parameters of the amplifier were experimentally determined: the small-signal gain per pass {approx}3.2, the linear gain {approx}0.031 cm{sup -1} with a nonuniformity of its distribution over the aperture within 15%, the stored energy of 0.16 - 0.21 J cm{sup -3}. The wavefront distortions in the zigzag laser-beam propagation through the active element of the amplifier did not exceed 0.4{lambda} ({lambda} = 0.63 {mu}m is the probing radiation wavelength).

  3. Sizing Power Components of an Electrically Driven Tail Cone Thruster and a Range Extender

    Science.gov (United States)

    Jansen, Ralph H.; Bowman, Cheryl; Jankovsky, Amy

    2016-01-01

    The aeronautics industry has been challenged on many fronts to increase efficiency, reduce emissions, and decrease dependency on carbon-based fuels. This paper provides an overview of the turboelectric and hybrid electric technologies being developed under NASA's Advanced Air Transportation Technology (AATT) Project and discusses how these technologies can impact vehicle design. The discussion includes an overview of key hybrid electric studies and technology investments, the approach to making informed investment decisions based on key performance parameters and mission studies, and the power system architectures for two candidate aircraft. Finally, the power components for a single-aisle turboelectric aircraft with an electrically driven tail cone thruster and for a hybrid-electric nine-passenger aircraft with a range extender are parametrically sized, and the sensitivity of these components to key parameters is presented.

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

    Energy Technology Data Exchange (ETDEWEB)

    Masherov, P. E.; Riaby, V. A., E-mail: riaby2001@yahoo.com [Research Institute of Applied Mechanics and Electrodynamics of the Moscow Aviation Institute (National Research University), Moscow (Russian Federation); Godyak, V. A. [Electrical Engineering and Computer Science Department, University of Michigan, Ann Arbor, Michigan 48109, USA and RF Plasma Consulting, Brookline, Massachusetts 02446 (United States)

    2016-02-15

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

  5. Hall Effect Thruster for High Power Solar Electric Propulsion Technology Demonstration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop a flight version of a high power Hall Effect thruster. While numerous high power Hall Effect thrusters have been demonstrated in the...

  6. Laser characterization of electric field oscillations in the Hall thruster breathing mode

    Science.gov (United States)

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

    2016-10-01

    Hall thrusters are a mature technology for space propulsion applications that exhibit a wide array of dynamic behavior, including plasma waves, instabilities and turbulence. One common low frequency (10-50 kHz) discharge current oscillation is the breathing mode, a cycle of neutral propellant injection, strong ionization, and ion acceleration by a steep potential gradient. A time-resolved laser-induced fluorescence diagnostic non-intrusively captures this propagating ionization front in the channel of a commercial BHT-600 Hall thruster manufactured by Busek Co. Measurements of ion velocity and relative ion density (using the 5 d[ 4 ] 7 / 2 - 6 p[ 3 ] 5 / 2 Xe II transition at 834.95 nm, vacuum) reveal a dynamic electric field structure traversing the channel throughout the breathing mode cycle. This work is sponsored by the U.S. Air Force Office of Scientific Research, with Dr. M. Birkan as program manager. C.Y. acknowledges support from the DOE NSSA Stewardship Science Graduate Fellowship under contract DE-FC52-08NA28752.

  7. Nanosecond Repetitively Pulsed Discharges in Air at Atmospheric Pressure -- Experiment and Theory of Regime Transitions

    Science.gov (United States)

    Pai, David; Lacoste, Deanna; Laux, Christophe

    2009-10-01

    In atmospheric pressure air preheated from 300 to 1000 K, the Nanosecond Repetitively Pulsed (NRP) method has been used to generate corona, glow, and spark discharges. Experiments have been performed to determine the parameter space (applied voltage, pulse repetition frequency, ambient gas temperature, and inter-electrode gap distance) of each discharge regime. Notably, there is a minimum gap distance for the existence of the glow regime that increases with decreasing gas temperature. A theory is developed to describe the Corona-to-Glow (C-G) and Glow-to-Spark (G-S) transitions for NRP discharges. The C-G transition is shown to depend on the Avalanche-to-Streamer Transition (AST) as well as the electric field strength in the positive column. The G-S transition is due to the thermal ionization instability. The minimum gap distance for the existence of the glow regime can be understood by considering that the applied voltage of the AST must be lower than that of the thermal ionization instability. This is a previously unknown criterion for generating glow discharges, as it does not correspond to the Paschen minimum or to the Meek-Raether criterion.

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

    Science.gov (United States)

    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. Overview of the Development of the Solar Electric Propulsion Technology Demonstration Mission 12.5-kW Hall Thruster

    Science.gov (United States)

    Kamhawi, Hani; Huang, Wensheng; Haag, Thomas; Yim, John; Chang, Li; Clayman, Lauren; Herman, Daniel; Shastry, Rohit; Thomas, Robert; Verhey, Timothy; Griffith, Christopher; Myers, James; Williams, George; Mikellides, Ioannis; Hofer, Richard; Polk, James; Goebel, Dan

    2014-01-01

    NASA is developing mission concepts for a solar electric propulsion technology demonstration mission. A number of mission concepts are being evaluated including ambitious missions to near Earth objects. The demonstration of a high-power solar electric propulsion capability is one of the objectives of the candidate missions under consideration. In support of NASA's exploration goals, a number of projects are developing extensible technologies to support NASA's near and long term mission needs. Specifically, the Space Technology Mission Directorate Solar Electric Propulsion Technology Demonstration Mission project is funding the development of a 12.5-kilowatt magnetically shielded Hall thruster system to support future NASA missions. This paper presents the design attributes of the thruster that was collaboratively developed by the NASA Glenn Research Center and the Jet Propulsion Laboratory. The paper provides an overview of the magnetic, plasma, thermal, and structural modeling activities that were carried out in support of the thruster design. The paper also summarizes the results of the functional tests that have been carried out to date. The planned thruster performance, plasma diagnostics (internal and in the plume), thermal, wear, and mechanical tests are outlined.

  10. Solid State Digital Propulsion "Cluster Thrusters" For Small Satellites Using High Performance Electrically Controlled Extinguishable Solid Propellants (ECESP)

    Science.gov (United States)

    Sawka, Wayne N.; Katzakian, Arthur; Grix, Charles

    2005-01-01

    Electrically controlled extinguishable solid propellants (ESCSP) are capable of multiple ignitions, extinguishments and throttle control by the application of electrical power. Both core and end burning no moving parts ECESP grains/motors to three inches in diameter have now been tested. Ongoing research has led to a newer family of even higher performance ECESP providing up to 10% higher performance, manufacturing ease, and significantly higher electrical conduction. The high conductivity was not found to be desirable for larger motors; however it is ideal for downward scaling to micro and pico- propulsion applications with a web thickness of less than 0.125 inch/ diameter. As a solid solution propellant, this ECESP is molecularly uniform, having no granular structure. Because of this homogeneity and workable viscosity it can be directly cast into thin layers or vacuum cast into complex geometries. Both coaxial and grain stacks have been demonstrated. Combining individual propellant coaxial grains and/or grain stacks together form three-dimensional arrays yield modular cluster thrusters. Adoption of fabless manufacturing methods and standards from the electronics industry will provide custom, highly reproducible micro-propulsion arrays and clusters at low costs. These stack and cluster thruster designs provide a small footprint saving spacecraft surface area for solar panels and/or experiments. The simplicity of these thrusters will enable their broad use on micro-pico satellites for primary propulsion, ACS and formation flying applications. Larger spacecraft may find uses for ECESP thrusters on extended booms, on-orbit refueling, pneumatic actuators, and gas generators.

  11. Low Mass Electromagnetic Plasmoid Thruster with Integrated PPU Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Electromagnetic Plasmoid Thruster (EMPT) is a revolutionary electric propulsion thruster and power processing (PPU) system that will allow a dramatic decrease...

  12. Hall Effect Thruster for High Power Solar Electric Propulsion Technology Demonstration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In Phase I Busek matured the design of an existing 15-kW laboratory thruster. Magnetic modeling was performed to generate a circuit incorporating magnetic shielding....

  13. Pulsed Electrogasdynamic Thruster for Attitude Control and Orbit Maneuver Project

    Data.gov (United States)

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

  14. Formation of the active medium in high-power repetitively pulsed gas lasers pumped by an electron-beam-controlled discharge

    Science.gov (United States)

    Bulaev, V. D.; Lysenko, S. L.

    2015-07-01

    A high-power repetitively pulsed e-beam-controlled discharge CO2 laser is simulated numerically; the simulation results are compared with experimental data. Optimal sizes and design of electrodes and configuration of the external magnetic field are found, which allow one to introduce no less than 90% electric pump energy into a specified volume of the active medium, including the active volume of a laser with an aperture of 110 × 110 cm. The results obtained can also be used to design other types of highpower gas lasers.

  15. The role of molecular vibration in nanosecond repetitively pulsed discharges and in DBDs in hydrogen plasmas

    Science.gov (United States)

    Colonna, G.; D'Ammando, G.; Pietanza, L. D.

    2016-10-01

    A self-consistent state-to-state model of pure hydrogen has been used to investigate the development of nanosecond repetitively pulsed discharges and dielectric barrier discharges, the latter coupling the kinetic model with an equation for the circuit, thus mimicking an insulated electrode with an external capacitance. Vibrationally excited states play a fundamental role, affecting the degrees of dissociation and ionization, as well as internal and free-electron distributions.

  16. Repetitively pulsed UV radiation source based on a run-away electron preionised diffuse discharge in nitrogen

    Science.gov (United States)

    Baksht, E. Kh; Burachenko, A. G.; Lomaev, M. I.; Panchenko, A. N.; Tarasenko, V. F.

    2015-04-01

    An extended repetitively pulsed source of spontaneous UV radiation is fabricated, which may also be used for producing laser radiation. Voltage pulses with an incident wave amplitude of up to 30 kV, a half-amplitude duration of ~4 ns and a rise time of ~2.5 ns are applied to a gap with a nonuniform electric field. For an excitation region length of 35 cm and a nitrogen pressure of 30 - 760 Torr, a diffusive discharge up to a pulse repetition rate of 2 kHz is produced without using an additional system for gap preionisation. An investigation is made of the plasma of the run-away electron preionised diffuse discharge. Using a CCD camera it is found that the dense diffused plasma fills the gap in a time shorter than 1 ns. X-ray radiation is recorded from behind the foil anode throughout the pressure range under study; a supershort avalanche electron beam is recorded by the collector electrode at pressures below 100 Torr.

  17. Transitions between corona, glow, and spark regimes of nanosecond repetitively pulsed discharges in air at atmospheric pressure

    Science.gov (United States)

    Pai, David Z.; Lacoste, Deanna A.; Laux, Christophe O.

    2010-05-01

    In atmospheric pressure air preheated from 300 to 1000 K, the nanosecond repetitively pulsed (NRP) method has been used to generate corona, glow, and spark discharges. Experiments have been performed to determine the parameter space (applied voltage, pulse repetition frequency, ambient gas temperature, and interelectrode gap distance) of each discharge regime. In particular, the experimental conditions necessary for the glow regime of NRP discharges have been determined, with the notable result that there exists a minimum and maximum gap distance for its existence at a given ambient gas temperature. The minimum gap distance increases with decreasing gas temperature, whereas the maximum does not vary appreciably. To explain the experimental results, an analytical model is developed to explain the corona-to-glow (C-G) and glow-to-spark (G-S) transitions. The C-G transition is analyzed in terms of the avalanche-to-streamer transition and the breakdown field during the conduction phase following the establishment of a conducting channel across the discharge gap. The G-S transition is determined by the thermal ionization instability, and we show analytically that this transition occurs at a certain reduced electric field for the NRP discharges studied here. This model shows that the electrode geometry plays an important role in the existence of the NRP glow regime at a given gas temperature. We derive a criterion for the existence of the NRP glow regime as a function of the ambient gas temperature, pulse repetition frequency, electrode radius of curvature, and interelectrode gap distance.

  18. Repetitively pulsed UV radiation source based on a run-away electron preionised diffuse discharge in nitrogen

    Energy Technology Data Exchange (ETDEWEB)

    Baksht, E Kh; Burachenko, A G; Lomaev, M I; Panchenko, A N; Tarasenko, V F [Institute of High Current Electronics, Siberian Branch, Russian Academy of Sciences, Tomsk (Russian Federation)

    2015-04-30

    An extended repetitively pulsed source of spontaneous UV radiation is fabricated, which may also be used for producing laser radiation. Voltage pulses with an incident wave amplitude of up to 30 kV, a half-amplitude duration of ∼4 ns and a rise time of ∼2.5 ns are applied to a gap with a nonuniform electric field. For an excitation region length of 35 cm and a nitrogen pressure of 30 – 760 Torr, a diffusive discharge up to a pulse repetition rate of 2 kHz is produced without using an additional system for gap preionisation. An investigation is made of the plasma of the run-away electron preionised diffuse discharge. Using a CCD camera it is found that the dense diffused plasma fills the gap in a time shorter than 1 ns. X-ray radiation is recorded from behind the foil anode throughout the pressure range under study; a supershort avalanche electron beam is recorded by the collector electrode at pressures below 100 Torr. (laser applications and other topics in quantum electronics)

  19. Investigation of Fe:ZnSe laser in pulsed and repetitively pulsed regimes

    Energy Technology Data Exchange (ETDEWEB)

    Velikanov, S D; Zaretskiy, N A; Zotov, E A; Maneshkin, A A; Chuvatkin, R S; Yutkin, I M [Russian Federal Nuclear Center ' All-Russian Research Institute of Experimental Physics' , Sarov, Nizhnii Novgorod region (Russian Federation); Kozlovsky, V I; Korostelin, Yu V; Krokhin, O N; Podmar' kov, Yu P; Savinova, S A; Skasyrsky, Ya K; Frolov, M P [P N Lebedev Physics Institute, Russian Academy of Sciences, Moscow (Russian Federation)

    2015-01-31

    The characteristics of a Fe:ZnSe laser pumped by a single-pulse free-running Er : YAG laser and a repetitively pulsed HF laser are presented. An output energy of 4.9 J is achieved in the case of liquid-nitrogen cooling of the Fe{sup 2+}:ZnSe active laser element longitudinally pumped by an Er:YAG laser with a pulse duration of 1 ms and an energy up to 15 J. The laser efficiency with respect to the absorbed energy is 47%. The output pulse energy at room temperature is 53 mJ. The decrease in the output energy is explained by a strong temperature dependence of the upper laser level lifetime and by pulsed heating of the active element. The temperature dependence of the upper laser level lifetime is used to determine the pump parameters needed to achieve high pulse energies at room temperature. Stable repetitively-pulsed operation of the Fe{sup 2+}:ZnSe laser at room temperature with an average power of 2.4 W and a maximum pulse energy of 14 mJ is achieved upon pumping by a 1-s train of 100-ns HF laser pulses with a repetition rate of 200 Hz. (lasers)

  20. Investigation of Fe:ZnSe laser in pulsed and repetitively pulsed regimes

    Science.gov (United States)

    Velikanov, S. D.; Zaretskiy, N. A.; Zotov, E. A.; Kozlovsky, V. I.; Korostelin, Yu V.; Krokhin, O. N.; Maneshkin, A. A.; Podmar'kov, Yu P.; Savinova, S. A.; Skasyrsky, Ya K.; Frolov, M. P.; Chuvatkin, R. S.; Yutkin, I. M.

    2015-01-01

    The characteristics of a Fe:ZnSe laser pumped by a single-pulse free-running Er : YAG laser and a repetitively pulsed HF laser are presented. An output energy of 4.9 J is achieved in the case of liquid-nitrogen cooling of the Fe2+:ZnSe active laser element longitudinally pumped by an Er:YAG laser with a pulse duration of 1 ms and an energy up to 15 J. The laser efficiency with respect to the absorbed energy is 47%. The output pulse energy at room temperature is 53 mJ. The decrease in the output energy is explained by a strong temperature dependence of the upper laser level lifetime and by pulsed heating of the active element. The temperature dependence of the upper laser level lifetime is used to determine the pump parameters needed to achieve high pulse energies at room temperature. Stable repetitively-pulsed operation of the Fe2+:ZnSe laser at room temperature with an average power of 2.4 W and a maximum pulse energy of 14 mJ is achieved upon pumping by a 1-s train of 100-ns HF laser pulses with a repetition rate of 200 Hz.

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

    Science.gov (United States)

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

    1978-01-01

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

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

    Science.gov (United States)

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

    1978-01-01

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

  3. Repetitively pulsed electric laser acoustic studies. Volume 1. Final technical report, Jun 80-Jun 83

    Energy Technology Data Exchange (ETDEWEB)

    Ingard, K.U.; McMillan, C.F.

    1983-09-01

    This report summarizes a study of the acoustical characteristics of a closed loop duct system for pulsed lasers with emphasis on acoustic suppression technology. Several topics are considered involving wave propagation reflection and attenuation in a shock tube, in which pulse waves are generated, simulating those in a pulsed laser system. A detailed analysis of the design of parallel-baffle attenuators for suppression of acoustic waves is given, allowing for the contributions of the reflection transmitted and reverberant contributions to the sound pressure field in the optical cavity.

  4. Solid State MEMS Thrusters Using Electrically Controlled Extinguishable Solid Propellant Project

    Data.gov (United States)

    National Aeronautics and Space Administration — ET Materials, LLC developed the first ever electrically controlled extinguishable solid propellant (ECESP). The original propellant developed under Air Force SBIR...

  5. Electra: durable repetitively pulsed angularly multiplexed KrF laser system

    Science.gov (United States)

    Wolford, Matthew F.; Myers, Matthew C.; Giuliani, John L.; Sethian, John D.; Burns, Patrick M.; Hegeler, Frank; Jaynes, Reginald

    2008-02-01

    Electra is a repetitively pulsed, electron beam pumped Krypton Fluoride (KrF) laser at the Naval Research Laboratory that is developing the technologies that can meet the Inertial Fusion Energy (IFE) requirements for durability, efficiency, and cost. The technologies developed on Electra should be directly scalable to a full size fusion power plant beam line. As in a full size fusion power plant beam line, Electra is a multistage laser amplifier system which, consists of a commercial discharge laser (LPX 305i, Lambda Physik), 175 keV electron beam pumped (40 ns flat-top) preamplifier, and 530 keV (100 ns flat-top) main amplifier. Angular multiplexing is used in the optical layout to provide pulse length control and to maximize laser extraction from the amplifiers. Single shot yield of 452 J has been extracted from the initial shots of the Electra laser system using a relatively low energy preamplifier laser beam. In rep-rate burst of 5 Hz for durations of one second a total energy of 1.585 kJ (average 317 J/pulse) has been attained. Total energy of 2.5 kJ has been attained over a two second period. For comparison, the main amplifier of Electra in oscillator mode has demonstrated at 2.5 Hz rep-rate average laser yield of 270 J over a 2 hour period.

  6. Low Mass Electromagnetic Plasmoid Thruster with Integrated PPU Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Electromagnetic Plasmoid Thruster (EMPT) is a revolutionary electric propulsion thruster and power processing (PPU) system that will allow a dramatic decrease in...

  7. Temperature Gradient in Hall Thrusters

    Energy Technology Data Exchange (ETDEWEB)

    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.

  8. Long Life 600W Hall Thruster System for Radioisotope Electric Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Radioisotope Electric Propulsion (REP) offers the prospect for a variety of new science missions by enabling use of Hall Effect propulsion in the outer solar system,...

  9. Transitions between corona, glow, and spark regimes of nanosecond repetitively pulsed discharges in air at atmospheric pressure

    OpenAIRE

    Pai, David,; Lacoste, Deanna,; Laux, C.

    2010-01-01

    International audience; In atmospheric pressure air preheated from 300 to 1000 K, the nanosecond repetitively pulsed (NRP) method has been used to generate corona, glow, and spark discharges. Experiments have been performed to determine the parameter space (applied voltage, pulse repetition frequency, ambient gas temperature, and interelectrode gap distance) of each discharge regime. In particular, the experimental conditions necessary for the glow regime of NRP discharges have been determine...

  10. Absolute OH Number Density Measurements in Lean Fuel-Air Mixtures Excited by a Repetitively Pulsed Nanosecond Discharge

    Science.gov (United States)

    2013-01-01

    discharge filaments and near the electrode edges [9]. Instead of using absorption measurement, an atmospheric pressure flame generated by a Hencken...DuPont) is placed between each electrode and the channel wall, to reduce air gaps and prevent corona discharge outside the cell. In the present work...1 Absolute OH Number Density Measurements in Lean Fuel-Air Mixtures Excited by a Repetitively Pulsed Nanosecond Discharge Zhiyao Yin, Campbell D

  11. A Microwave Thruster for Spacecraft Propulsion

    Energy Technology Data Exchange (ETDEWEB)

    Chiravalle, Vincent P [Los Alamos National Laboratory

    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.

  12. PERIPHERAL APPLICATION OF REPETITIVE PULSE MAGNETIC STIMULATION ON JOINT CONTRACTURE FOR MOBILITY RESTORATION: CONTROLLED RANDOMIZED STUDY

    Directory of Open Access Journals (Sweden)

    Efthimios J. Kouloulas

    2016-10-01

    Full Text Available Background: Joint contracture is a limitation in the passive or active range of motion (ROM of a joint, where in addition to the mobility limiting factor the pain is also present. Repetitive pulsed Magnetic Stimulation (rPMS appears to be an effective, non-invasive and safety solution for treating this condition. Therefore aim of this study was to evaluate the effect of rPMS in treating joint contracture. Methods: 30 subjects with joint contracture in the knee were enrolled in this study and divided respectively into Treatment and Control group. The treatment group were delivered with rPMS therapy. The control group was delivered with conventional physiotherapy method (ultrasound. The primary outcome measurements were: 1. Mobility evaluation by goniometry (ROM in degrees while performing flexion and Patient Functional Assessment Questionnaire (PFAQ for ability to perform Activities of Daily Living (ADL and 2. Pain evaluation by 10-point Visual Analog Scale (VAS for pain perception. Absence of adverse events was set as a secondary measure. Results: The results of the study show statistical difference (p<0.05 between the levels of improvement of all studied parameters while comparing between both groups. The results suggest greater immobility restoration and pain relieving effect of the rPMS in comparison to conventional physiotherapy method. Conclusion: rPMS an effective and safe non-invasive method for mobility restoration and pain relief in case of joint contractures. This study suggests the method as beneficial and quality of life ameliorating among patients suffering from immobilized joints accompanied by pain.

  13. Nanosecond repetitively pulsed discharges in air at atmospheric pressure—the spark regime

    Science.gov (United States)

    Pai, David Z.; Lacoste, Deanna A.; Laux, Christophe O.

    2010-12-01

    Nanosecond repetitively pulsed (NRP) spark discharges have been studied in atmospheric pressure air preheated to 1000 K. Measurements of spark initiation and stability, plasma dynamics, gas temperature and current-voltage characteristics of the spark regime are presented. Using 10 ns pulses applied repetitively at 30 kHz, we find that 2-400 pulses are required to initiate the spark, depending on the applied voltage. Furthermore, about 30-50 pulses are required for the spark discharge to reach steady state, following initiation. Based on space- and time-resolved optical emission spectroscopy, the spark discharge in steady state is found to ignite homogeneously in the discharge gap, without evidence of an initial streamer. Using measured emission from the N2 (C-B) 0-0 band, it is found that the gas temperature rises by several thousand Kelvin in the span of about 30 ns following the application of the high-voltage pulse. Current-voltage measurements show that up to 20-40 A of conduction current is generated, which corresponds to an electron number density of up to 1015 cm-3 towards the end of the high-voltage pulse. The discharge dynamics, gas temperature and electron number density are consistent with a streamer-less spark that develops homogeneously through avalanche ionization in volume. This occurs because the pre-ionization electron number density of about 1011 cm-3 produced by the high frequency train of pulses is above the critical density for streamer-less discharge development, which is shown to be about 108 cm-3.

  14. Hall-Effect Thruster Modifications for Dual-Mode Electric Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The integrated NASA/DoD electric propulsion objectives are for a specific mass less than 3 kg/kW while demonstrating a throttlable thrust-to-power ratio of 100:1 at...

  15. Electrodeless plasma thrusters for spacecraft: A review

    Science.gov (United States)

    Bathgate, S. N.; Bilek, M. M. M.; McKenzie, D. R.

    2017-08-01

    The physics of electrodeless electric thrusters that use directed plasma to propel spacecraft without employing electrodes subject to plasma erosion is reviewed. Electrodeless plasma thrusters are potentially more durable than presently deployed thrusters that use electrodes such as gridded ion, Hall thrusters, arcjets and resistojets. Like other plasma thrusters, electrodeless thrusters have the advantage of reduced fuel mass compared to chemical thrusters that produce the same thrust. The status of electrodeless plasma thrusters that could be used in communications satellites and in spacecraft for interplanetary missions is examined. Electrodeless thrusters under development or planned for deployment include devices that use a rotating magnetic field; devices that use a rotating electric field; pulsed inductive devices that exploit the Lorentz force on an induced current loop in a plasma; devices that use radiofrequency fields to heat plasmas and have magnetic nozzles to accelerate the hot plasma and other devices that exploit the Lorentz force. Using metrics of specific impulse and thrust efficiency, we find that the most promising designs are those that use Lorentz forces directly to expel plasma and those that use magnetic nozzles to accelerate plasma.

  16. Drilling and cutting of thin metal plates in water with radiation of a repetitively pulsed Nd : YAG laser

    Science.gov (United States)

    Glova, A. F.; Lysikov, A. Yu

    2011-10-01

    The conditions of drilling and cutting of 0.15-mm-thick titanium and stainless steel plates in water with the radiation of a repetitively pulsed Nd : YAG laser having the mean power up to 30 W are studied experimentally in the absence of water and gas jets. Dependences of the maximal cutting speed in water on the radiation power are obtained, the cutting efficiency is determined, and the comparison with the conditions of drilling and cutting of plates in air is carried out.

  17. Helical plasma thruster

    Energy Technology Data Exchange (ETDEWEB)

    Beklemishev, A. D., E-mail: bekl@bk.ru [Budker Institute of Nuclear Physics SB RAS, Novosibirsk (Russian Federation)

    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. Thruster Module

    Science.gov (United States)

    Andersson, G.

    2015-09-01

    The thruster module described in this paper provides a low but controlled acceleration in a mission which would normally be labelled “microgravity”. The first mission was Cryofenix, where tanks containing liquid hydrogen were used in the experiment. The experiment utilizing the low acceleration is using liquids and requires a precise acceleration profile throughout the mission. Acceleration obtained by payload rotation is not feasible due to that the transversal forces required to change the acceleration will cause undesired liquid turbulence. In order to satisfy the experiment requirements a thruster module was developed by SSC for the Cryofenix mission funded by CNES. The Cryofenix mission had a payload weight of 380 kg and an apogee of about 260 km. The module produces a controlled thrust in flight direction by means of a cold gas system.

  19. A comparison of inflection point and floating point emissive probe techniques for electric potential measurements in a Hall thruster plasma

    Science.gov (United States)

    Sheehan, J. P.; Raitses, Yevgeny; Hershkowitz, Noah; Fisch, Nathaniel

    2010-11-01

    Theory suggests that when increasing the electron emission of an emissive probe the floating potential will saturate ˜Te/e below the plasma potential. This can introduce significant errors in plasma potential measurements in Hall thrusters where Te> 10 eV. The method of determining the plasma potential from the inflection point of emissive IV traces in the limit of zero emission may give a more accurate measurement of the plasma potential. The two methods are compared in a Hall thruster where ne˜10^11 cm-3, Te˜20 eV, and ion flows are significant. The results can be generalized to other types of plasmas.

  20. Thermal Management of Superconducting Electromagnets in VASIMR Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Future manned space exploration missions will require high power electric propulsion. VASIMR thrusters are the most attractive option because they offer short...

  1. Control Valve for Miniature Xenon Ion Thruster Project

    Data.gov (United States)

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

  2. Effects of pulse-to-pulse residual species on discharges in repetitively pulsed discharges through packed bed reactors

    Science.gov (United States)

    Kruszelnicki, Juliusz; Engeling, Kenneth W.; Foster, John E.; Kushner, Mark J.

    2016-09-01

    Atmospheric pressure dielectric barrier discharges (DBDs) sustained in packed bed reactors (PBRs) are being investigated for conversion of toxic and waste gases, and CO2 removal. These discharges are repetitively pulsed having varying flow rates and internal geometries, which results in species from the prior pulse still being in the discharge zone at the time the following discharge pulse occurs. A non-negligible residual plasma density remains, which effectively acts as preionization. This residual charge changes the discharge properties of subsequent pulses, and may impact important PBR properties such as chemical selectivity. Similarly, the residual neutral reactive species produced during earlier pulses will impact the reaction rates on subsequent pulses. We report on results of a computational investigation of a 2D PBR using the plasma hydrodynamics simulator nonPDPSIM. Results will be discussed for air flowing though an array of dielectric rods at atmospheric pressure. The effects of inter-pulse residual species on PBR discharges will be quantified. Means of controlling the presence of residual species in the reactor through gas flow rate, pulse repetition, pulse width and geometry will be described. Comparisons will be made to experiments. Work supported by US DOE Office of Fusion Energy Science and the National Science Foundation.

  3. Q-thruster Breadboard Campaign Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Q-thruster technology is a mission enabling form of electric propulsion and is already being traded by NASA's Concept Architecture Team (CAT) & Human Space...

  4. Iodine Hall Thruster for Space Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek Co. Inc. proposes to develop a high power (high thrust) electric propulsion system featuring an iodine fueled Hall Effect Thruster (HET). The system to be...

  5. Additive Manufacturing of Ion Thruster Optics Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Plasma Controls will manufacture and test a set of ion optics for electric propulsion ion thrusters using additive manufacturing technology, also known as 3D...

  6. Arcjet space thrusters

    Science.gov (United States)

    Keefer, Dennis; Rhodes, Robert

    1993-01-01

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

  7. Pseudospectral Model for Hybrid PIC Hall-effect Thruster Simulation

    Science.gov (United States)

    2015-07-01

    1149. 8Goebel, D. M. and Katz, I., Fundamentals of Electric Propulsion : Ion and Hall Thrusters, John Wiley & Sons, Inc., 2008. 9Martin, R., J.W., K...Bilyeu, D., and Tran, J., “Dynamic Particle Weight Remapping in Hybrid PIC Hall -effect Thruster Simulation,” 34th Int. Electric Propulsion Conf...Paper 3. DATES COVERED (From - To) July 2015-July 2015 4. TITLE AND SUBTITLE Pseudospectral model for hybrid PIC Hall -effect thruster simulationect

  8. Electron dynamics in Hall thruster

    Science.gov (United States)

    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.

  9. Miniature Bipolar Electrostatic Ion Thruster

    Science.gov (United States)

    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.

  10. Enhanced performance of a repetitively pulsed 130 mJ KrF laser with improved pre-ionization parameters

    Indian Academy of Sciences (India)

    N S Benerji; A Singh; N Varshnay; Bijendra Singh

    2014-01-01

    Studies related to the effect of pre-ionizer on laser output energy of a repetitively pulsed KrF laser are presented. The dependence of laser output energy, spectral width and beam spot homogeneity on pre-ionization parameters, namely its current and voltage rise time are reported here. Here, effectiveness of pre-ionization is optimized by improving pre-ionization current and rise time of the pump pulse of the automatic UV pre-ionizer KrF laser. It is observed that by increasing pre-ionization current from 6 kA to 10.6 kA, the output energy increases by about 30% (from 100 to 130 mJ). It is also observed that the emission spectral width reduces by almost 60% by increasing the pre-ionization current. Regular homogeneous and well-developed beam spot (nearly Hat-Top profile) was achieved under these optimized conditions.

  11. Wavelength dependence of repetitive-pulse laser-induced damage threshold in beta-BaB2O4.

    Science.gov (United States)

    Kouta, H

    1999-01-20

    The dependence on wavelength of repetitive-pulse (10 Hz, 8-10 ns) laser-induced damage on beta barium metaborate (BBO) has been investigated. The thresholds of dielectric breakdown in bulk crystal have been found to be 0.3 GW/cm(2) at 266 nm, 0.9 GW/cm(2) at 355 nm, 2.3 GW/cm(2) at 532 nm, and 4.5 GW/cm(2) at 1064 nm. Results indicate two-photon absorption at 266 and 355 nm, which helps to produce an avalanche effect that causes breakdown at each of the four wavelengths tested. Neither the BBO refractive indices nor the absorption spectrum change until breakdown occurs.

  12. The influence of repetitively pulsed plasma immersion low energy ion implantation on TiN coating formation and properties

    Science.gov (United States)

    Sivin, D. O.; Ananin, P. S.; Dektyarev, S. V.; Ryabchikov, A. I.; Shevelev, A. E.

    2017-05-01

    Application of high frequency short pulse plasma immersion low energy ion implantation for titanium nitride coating deposition using vacuum arc metal plasma and hot-cathode gas-discharge plasma on R6M5 alloy was investigated. Implementation of negative repetitively pulsed bias with bias amplitude 2 kV, pulse duration 5 μs and pulse frequency 105 Hz leads to 6.2-fold decrease of vacuum arc macroparticle surface density for macroparticles with diameter less than 0.5 μm. Ion sputtering due coating deposition reduces the production rate approximately by 30%. It was found that with bias amplitude range from 1.1 to 1.4 kV and pulse duration 5 μs yields to formation of coatings with local hardness up to 40 GPa. This paper presents the results of experimental studies of adhesion strength, tribological properties and surface morphology of deposited TiN coatings.

  13. Mechanical design of SERT 2 thruster system

    Science.gov (United States)

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

    1972-01-01

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

  14. Characteristics of the evolution of a plasma generated by radiation from CW and repetitively pulsed CO2 lasers in different gases

    Science.gov (United States)

    Kanevskii, M. F.; Stepanova, M. A.

    1990-06-01

    The interaction between high-power CW and repetitively pulsed CO2 laser radiation and a low-threshold optical-breakdown plasma near a metal surface is investigated. The characteristics of the breakdown plasma are examined as functions of the experimental conditions. A qualitative analysis of the results obtained was performed using a simple one-dimensional model for laser combustion waves.

  15. A pulse-compression-ring circuit for high-efficiency electric propulsion.

    Science.gov (United States)

    Owens, Thomas L

    2008-03-01

    A highly efficient, highly reliable pulsed-power system has been developed for use in high power, repetitively pulsed inductive plasma thrusters. The pulsed inductive thruster ejects plasma propellant at a high velocity using a Lorentz force developed through inductive coupling to the plasma. Having greatly increased propellant-utilization efficiency compared to chemical rockets, this type of electric propulsion system may one day propel spacecraft on long-duration deep-space missions. High system reliability and electrical efficiency are extremely important for these extended missions. In the prototype pulsed-power system described here, exceptional reliability is achieved using a pulse-compression circuit driven by both active solid-state switching and passive magnetic switching. High efficiency is achieved using a novel ring architecture that recovers unused energy in a pulse-compression system with minimal circuit loss after each impulse. As an added benefit, voltage reversal is eliminated in the ring topology, resulting in long lifetimes for energy-storage capacitors. System tests were performed using an adjustable inductive load at a voltage level of 3.3 kV, a peak current of 20 kA, and a current switching rate of 15 kA/micros.

  16. Influence of air flow parameters on nanosecond repetitively pulsed discharges in a pin-annular electrode configuration

    KAUST Repository

    Heitz, Sylvain A

    2016-03-16

    The effect of various air flow parameters on the plasma regimes of nanosecond repetitively pulsed (NRP) discharges is investigated at atmospheric pressure. The two electrodes are in a pin-annular configuration, transverse to the mean flow. The voltage pulses have amplitudes up to 15 kV, a duration of 10 ns and a repetition frequency ranging from 15 to 30 kHz. The NRP corona to NRP spark (C-S) regime transition and the NRP spark to NRP corona (S-C) regime transition are investigated for different steady and harmonically oscillating flows. First, the strong effect of a transverse flow on the C-S and S-C transitions, as reported in previous studies, is verified. Second, it is shown that the azimuthal flow imparted by a swirler does not affect the regime transition voltages. Finally, the influence of low frequency harmonic oscillations of the air flow, generated by a loudspeaker, is studied. A strong effect of frequency and amplitude of the incoming flow modulation on the NRP plasma regime is observed. Results are interpreted based on the cumulative effect of the NRP discharges and an analysis of the residence times of fluid particles in the inter-electrode region. © 2016 IOP Publishing Ltd.

  17. Investigation of Gas Heating by Nanosecond Repetitively Pulsed Glow Discharges Used for Actuation of a Laminar Methane-Air Flame

    KAUST Repository

    Lacoste, Deanna A.

    2017-05-24

    This paper reports on the quantification of the heating induced by nanosecond repetitively pulsed (NRP) glow discharges on a lean premixed methane-air flame. The flame, obtained at room temperature and atmospheric pressure, has an M-shape morphology. The equivalence ratio is 0.95 and the thermal power released by the flame is 113 W. The NRP glow discharges are produced by high voltage pulses of 10 ns duration, 7 kV amplitude, applied at a repetition frequency of 10 kHz. The average power of the plasma, determined from current and voltage measurements, is 1 W, i.e. about 0.9 % of the thermal power of the flame. Broadband vibrational coherent anti-Stokes Raman spectroscopy of nitrogen is used to determine the temperature of the flame with and without plasma enhancement. The temperature evolution in the flame area shows that the thermal impact of NRP glow discharges is in the uncertainty range of the technique, i.e., +/- 40 K.

  18. Recent developments in high-resolution optical diagnostics of repetitively pulsed laser-target effects

    Science.gov (United States)

    Hugenschmidt, Manfred; Althaus, Marion

    1995-05-01

    High energy densities, as required both in research and in industry, are achieved by the use of lasers. Extremely highpower densities are obtained in the pulsed mode with short microsecond(s) -, ns-, or even ultrashort ps- to fs- pulses. The interaction of such powerful laser pulses with any type of solid state, liquid or gaseous materials is then causing rapidly developing, nonstationary, optically nonlinear processes. Experimental investigations of these effects are therefore requiring special measuring techniques with high spatial and temporal resolution. Optical and optronical methods have proven to be particularly useful. Methods based on laser diagnostics, including high speed photography, cinematography, speckle techniques, holography, videography, infrared techniques or arbitrary combinations of these, are therefore considered to be important tools in these laser effect studies. The investigations reported in the present paper are referring to carbon dioxide-laser effects in intensity ranges which are useful for many industrial applications, such as for example in the field of material processing. Basic interest is actually in pulsed, plasma sustained laser target interaction phenomena which occur above critical threshold power densities, specific for each type of material. Surface induced, highly ionized absorption waves are then determining the energy transfer from the coherent laser radiation field towards the targets. The experiments at ISL were aimed at investigating plasma parameters and their influence on the energy transfer rates, by fast optical, electrical and optronical techniques, such as mentioned above. The results to be discussed refer to target effects, basically observed on optically transparent materials, subject to high average power pulsed carbon dioxide-laser radiation, with repetition rates of several tens to hundred pps at multi-MW/cm2 to GW/cm2 peak power densities and average power densities in the multi-kW/cm2-range.

  19. Plasma Diagnostic and Performance of a Permanent Magnet Hall Thruster

    CERN Document Server

    Ferreira, J L; Rego, I D S; Ferreira, I S; Ferreira, Jose Leonardo; Souza, Joao Henrique Campos De; Rego, Israel Da Silveira; Ferreira, Ivan Soares

    2004-01-01

    Electric propulsion is now a sucessfull method for primary propulsion of deep space long duration missions and for geosyncronous satellite attitude control. Closed Drift Plasma Thruster, so called Hall Thruster or SPT (stationary plasma thruster) were primarily conceived in USSR (the ancient Soviet Union) and now it is been developed by space agencies, space research institutes and industries in several countries such as France, USA, Israel, Russian Federation and Brazil. In this work, we show plasma characteristics and performance of a Hall Thruster designed with an innovative concept which uses an array of permanent magnets, instead of an eletromagnet, to produce a radial magnetic field inside its cylindrical plasma drift channel. Within this new concept, we expect to develop a Hall Thruster within power consuption that will scale up to small and medium size satellites. A plasma density and temperature space profiles inside and outside the thruster channel will be shown. Space plasma potential, ion temperat...

  20. Los Alamos NEP research in advanced plasma thrusters

    Science.gov (United States)

    Schoenberg, Kurt; Gerwin, Richard

    1991-01-01

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

  1. Laser generation of XeCl exciplex molecules in a longitudinal repetitively pulsed discharge in a Xe – CsCl mixture

    Energy Technology Data Exchange (ETDEWEB)

    Boichenko, A M [A M Prokhorov General Physics Institute, Russian Academy of Sciences, Moscow (Russian Federation); Klenovskii, M S [National Research Tomsk Polytechnic University, Tomsk (Russian Federation)

    2015-12-31

    By using the previously developed kinetic model, we have carried out simulations to study the possibility of laser generation of XeCl exciplex molecules in the working medium based on a mixture of Xe with CsCl vapours, excited by a longitudinal repetitively pulsed discharge. The formation mechanism of exciplex molecules in this mixture is fundamentally different from the formation mechanisms in the traditional mixtures of exciplex lasers. The conditions that make the laser generation possible are discussed. For these conditions, with allowance for available specific experimental conditions of the repetitively pulsed discharge excitation, we have obtained the calculated dependences of the power and efficiency of generation on the reflectivity of mirrors in a laser cavity. (active media)

  2. A repetitively pulsed xenon chloride excimer laser with all ferrite magnetic cores (AFMC) based all solid state exciter

    Science.gov (United States)

    Benerji, N. S.; Varshnay, N. K.; Ghodke, D. V.; Singh, A.

    2016-10-01

    Performance of repetitively pulsed xenon chloride excimer laser (λ~308 nm) with solid state pulser consisting of magnetic pulse compression circuit (MPC) using all ferrite magnetic cores (AFMC) is reported. Laser system suitable for 100 Hz operation with inbuilt pre-ionizer, compact gas circulation and cooling has been developed and presented. In this configuration, high voltage pulses of ~8 μs duration are compressed to ~100 ns by magnetic pulse compression circuit with overall compression factor of ~80. Pulse energy of ~18 J stored in the primary capacitor is transferred to the laser head with an efficiency of ~85% compared to ~70% that is normally achieved in such configurations using annealed met-glass core. This is a significant improvement of about 21%. Maximum output laser pulse energy of ~100 mJ was achieved at repetition rate of 100 Hz with a typical pulse to pulse energy stability of ±5% and laser pulse energy of 150 mJ was generated at low rep-rate of ~40 Hz. This exciter uses a low current and low voltage solid state switch (SCR) that replaces high voltage and high current switch i. e, thyratron completely. The use of solid state exciter in turn reduces electromagnetic interference (EMI) effects particularly in excimer lasers where high EMI is present due to high di/dt. The laser is focused on a thin copper sheet for generation of micro-hole and the SEM image of the generated micro hole shows the energy stability of the laser at high repetition rate operation. Nearly homogeneous, regular and well developed xenon chloride (XeCl) laser beam spot was achieved using the laser.

  3. Cylindrical geometry hall thruster

    Science.gov (United States)

    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.

  4. Conducting Wall Hall Thrusters

    Science.gov (United States)

    Goebel, Dan M.; Hofer, Richard R.; Mikellides, Ioannis G.; Katz, Ira; Polk, James E.; Dotson, Brandon

    2013-01-01

    A unique configuration of the magnetic field near the wall of Hall thrusters, called Magnetic Shielding, has recently demonstrated the ability to significantly reduce the erosion of the boron nitride (BN) walls and extend the life of Hall thrusters by orders of magnitude. The ability of magnetic shielding to minimize interactions between the plasma and the discharge chamber walls has for the first time enabled the replacement of insulating walls with conducting materials without loss in thruster performance. The boron nitride rings in the 6 kW H6 Hall thruster were replaced with graphite that self-biased to near the anode potential. The thruster efficiency remained over 60% (within two percent of the baseline BN configuration) with a small decrease in thrust and increase in Isp typical of magnetically shielded Hall thrusters. The graphite wall temperatures decreased significantly compared to both shielded and unshielded BN configurations, leading to the potential for higher power operation. Eliminating ceramic walls makes it simpler and less expensive to fabricate a thruster to survive launch loads, and the graphite discharge chamber radiates more efficiently which increases the power capability of the thruster compared to conventional Hall thruster designs.

  5. Oxygen-Methane Thruster

    Science.gov (United States)

    Pickens, Tim

    2012-01-01

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

  6. Conducting wall Hall thrusters in magnetic shielding and standard configurations

    Science.gov (United States)

    Grimaud, Lou; Mazouffre, Stéphane

    2017-07-01

    Traditional Hall thrusters are fitted with boron nitride dielectric discharge channels that confine the plasma discharge. Wall properties have significant effects on the performances and stability of the thrusters. In magnetically shielded thrusters, interactions between the plasma and the walls are greatly reduced, and the potential drop responsible for ion acceleration is situated outside the channel. This opens the way to the utilization of alternative materials for the discharge channel. In this work, graphite walls are compared to BN-SiO2 walls in the 200 W magnetically shielded ISCT200-MS and the unshielded ISCT200-US Hall thrusters. The magnetically shielded thruster shows no significant change in the discharge current mean value and oscillations, while the unshielded thruster's discharge current increases by 25% and becomes noticeably less stable. The electric field profile is also investigated through laser spectroscopy, and no significant difference is recorded between the ceramic and graphite cases for the shielded thruster. The unshielded thruster, on the other hand, has its acceleration region shifted 15% of the channel length downstream. Lastly, the plume profile is measured with planar probes fitted with guard rings. Once again the material wall has little influence on the plume characteristics in the shielded thruster, while the unshielded one is significantly affected.

  7. Advanced laboratory for testing plasma thrusters and Hall thruster measurement campaign

    Directory of Open Access Journals (Sweden)

    Szelecka Agnieszka

    2016-06-01

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

  8. A high power ion thruster for deep space missions

    Science.gov (United States)

    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.

  9. Three Phase Resonant DC Power Converter for Ion Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The new generation of, high performance electric propulsion missions will require high mass throughput and most likely the use of grided ion thruster equipped with...

  10. Diagnostics Systems for Permanent Hall Thrusters Development

    Science.gov (United States)

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

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

  11. Laser thruster

    Science.gov (United States)

    Jalufka, N. W.

    1989-01-01

    Laser propulsion can reduce fuel by 57 t to 105 t over chemical propulsion for a 144 t Lunar base, with no significant increase in trip time. Laser propulsion reduces trip time by a factor of 40 to 120 over nuclear electric propulsion and time in radiation belts by a factor of 100 to 1700. Either solar or nuclear driven laser diode arrays could produce multimegawatt beams, typically 3,700 t for a 235 MW laser system. Laser diode arrays have high payoff due to short wavelength (850 nm) and high diode efficiency (70 percent). A dry laser OTV of 8790 kg and 60 percent efficiency can transport a 144 t lunar base. Laser propulsion could carry both personnel and cargo safely to the lunar base.

  12. Experimental determination of plasma detachment from the diverging magnetic nozzle of the VASIMR VX-200 Electric Thruster

    Science.gov (United States)

    Olsen, Christopher; Squire, Jared; Longmier, Benjamin; Ballenger, Maxwell; Cassady, Leonard; Carter, Mark; Ilin, Andrew; Cloutier, Paul; Bering, Edgar; Giambusso, Matthew; Ad Astra Rocket Company Team; Rice University Collaboration; University of Houston Collaboration

    2011-10-01

    Theories of magnetized plasma detachment in an expanding magnetic field have been lacking detailed experimental evidence. Recent experiments using a 200 kW class electric rocket (VX-200), run at 100 kW using argon and a peak magnetic field of 2 T, produced ion energies greater than 100 eV with a flux of 2x1022 ions/s in a 150 m3 vacuum facility. Ion-neutral charge exchange effects were reduced and the resultant data show evidence of plasma detachment in a diverging magnetic field on a scale length of 2 m. The detachment is confirmed using multiple plasma diagnostics and magnetic nozzle topologies. Spatial maps of the data are compared to simulations from a particle detachment model, ParTraj, as well as MHD detachment theory. ParTraj, when compared to experiment, is shown to be more consistent in describing the data. Unless the MHD models are modified to incorporation two-fluid effects, single fluid MHD theory is inconsistent with the observations.

  13. Performance and flow characteristics of MHD seawater thruster

    Energy Technology Data Exchange (ETDEWEB)

    Doss, E.D.

    1990-01-01

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

  14. Low-Mass, Low-Power Hall Thruster System

    Science.gov (United States)

    Pote, Bruce

    2015-01-01

    NASA is developing an electric propulsion system capable of producing 20 mN thrust with input power up to 1,000 W and specific impulse ranging from 1,600 to 3,500 seconds. The key technical challenge is the target mass of 1 kg for the thruster and 2 kg for the power processing unit (PPU). In Phase I, Busek Company, Inc., developed an overall subsystem design for the thruster/cathode, PPU, and xenon feed system. This project demonstrated the feasibility of a low-mass power processing architecture that replaces four of the DC-DC converters of a typical PPU with a single multifunctional converter and a low-mass Hall thruster design employing permanent magnets. In Phase II, the team developed an engineering prototype model of its low-mass BHT-600 Hall thruster system, with the primary focus on the low-mass PPU and thruster. The goal was to develop an electric propulsion thruster with the appropriate specific impulse and propellant throughput to enable radioisotope electric propulsion (REP). This is important because REP offers the benefits of nuclear electric propulsion without the need for an excessively large spacecraft and power system.

  15. Single and Multi-Pulse Low-Energy Conical Theta Pinch Inductive Pulsed Plasma Thruster Performance

    Science.gov (United States)

    Hallock, Ashley K.; Martin, Adam; Polzin, Kurt; Kimberlin, Adam; Eskridge, Richard

    2013-01-01

    Fabricated and tested CTP IPPTs at cone angles of 20deg, 38deg, and 60deg, and performed direct single-pulse impulse bit measurements with continuous gas flow. Single pulse performance highest for 38deg angle with impulse bit of approx.1 mN-s for both argon and xenon. Estimated efficiencies low, but not unexpectedly so based on historical data trends and the direction of the force vector in the CTP. Capacitor charging system assembled to provide rapid recharging of capacitor bank, permitting repetition-rate operation. IPPT operated at repetition-rate of 5 Hz, at maximum average power of 2.5 kW, representing to our knowledge the highest average power for a repetitively-pulsed thruster. Average thrust in repetition-rate mode (at 5 kV, 75 sccm argon) was greater than simply multiplying the single-pulse impulse bit and the repetition rate.

  16. Thermal Modeling for Pulsed Inductive FRC Plasmoid Thrusters

    Science.gov (United States)

    Pfaff, Michael

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

  17. Oxygen-Methane Thruster Project

    Data.gov (United States)

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

  18. Iodine Hall Thruster

    Science.gov (United States)

    Szabo, James

    2015-01-01

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

  19. Global model of an iodine gridded plasma thruster

    Science.gov (United States)

    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.

  20. A collisionless plasma thruster plume expansion model

    Science.gov (United States)

    Merino, Mario; Cichocki, Filippo; Ahedo, Eduardo

    2015-06-01

    A two-fluid model of the unmagnetized, collisionless far region expansion of the plasma plume for gridded ion thrusters and Hall effect thrusters is presented. The model is integrated into two semi-analytical solutions valid in the hypersonic case. These solutions are discussed and compared against the results from the (exact) method of characteristics; the relative errors in density and velocity increase slowly axially and radially and are of the order of 10-2-10-3 in the cases studied. The plasma density, ion flux and ambipolar electric field are investigated. A sensitivity analysis of the problem parameters and initial conditions is carried out in order to characterize the far plume divergence angle in the range of interest for space electric propulsion. A qualitative discussion of the physics of the secondary plasma plume is also provided.

  1. Magnesium Hall Thruster

    Science.gov (United States)

    Szabo, James J.

    2015-01-01

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

  2. Performance Evaluation of the Prototype Model NEXT Ion Thruster

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    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.

  4. Liquid micro pulsed plasma thruster

    Directory of Open Access Journals (Sweden)

    Szelecka Agnieszka

    2015-06-01

    Full Text Available A new type of pulsed plasma thruster (PPT for small satellite propulsion is investigated, of which the most innovative aspect is the use of a non-volatile liquid propellant. The thruster is based on an open capillary design. The thruster achieved a thrust-to-power ratio above 45 μN/W, which constitutes a 5-fold improvement over the water-propelled pulsed plasma thruster, and which is also slightly above the performance of a similarly sized PPT with a solid propellant.

  5. Effect of Ambipolar Potential on the Propulsive Performance of the GDM Plasma Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Gasdynamic Mirror (GDM) thruster is an electric propulsion device, without electrodes, that will magnetically confine a plasma with such density and temperature...

  6. INTERACTION OF LASER RADIATION WITH MATTER. LASER PLASMA: Calculation of mass transfer in the remote cutting of metals by radiation of a high-power repetitively pulsed CO2 laser

    Science.gov (United States)

    Gladush, G. G.; Rodionov, N. B.

    2002-01-01

    The mechanism of remote cutting of steel plates by radiation of a high-power repetitively pulsed CO2 laser is theoretically studied. The models of melt removal by the gravity force and the recoil pressure of material vapour are proposed and the sufficient conditions for the initiation of cutting are determined. A numerical model of a thermally thin plate was employed to describe the cutting for large focal spots.

  7. Dynamic Particle Weight Remapping in Hybrid PIC Hall-effect Thruster Simulation

    Science.gov (United States)

    2015-05-01

    International Electric Propulsion Conference and 6th Nano-satellite Symposium Hyogo-Kobe, Japan July 410, 2015 Robert Martin∗ ERC Incorporated, Huntsville...Algorithms, . 8Koo, J. and Martin, R., Pseudospectral model for hybrid PIC Hall -eect thruster simulation, 34th Int. Electric Propul- sion Conf...Paper 3. DATES COVERED (From - To) May 2015-July 2015 4. TITLE AND SUBTITLE Dynamic Particle Weight Remapping in Hybrid PIC Hall -effect Thruster

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

    Science.gov (United States)

    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

  9. An approach to the parametric design of ion thrusters

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    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.

  11. Thermal Environmental Testing of NSTAR Engineering Model Ion Thrusters

    Science.gov (United States)

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

    1999-01-01

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

  12. Magnetic Field Tailored Annular Hall Thruster with Anode Layer

    Science.gov (United States)

    Lee, Seunghun; Kim, Holak; Kim, Junbum; Lim, Youbong; Choe, Wonho; Korea Institute of Materials Science Collaboration

    2016-09-01

    Plasma propulsion system is one of the key components for advanced missions of satellites as well as deep space exploration. A typical plasma propulsion system is Hall effect thruster that uses crossed electric and magnetic fields to ionize a propellant gas and to accelerate the ionized gas to generate momentum. In Hall thruster plasmas, magnetic field configuration is important due to the fact that electron confinement in the electromagnetic fields affects both plasma and ion beam characteristics as well as thruster performance parameters including thrust, specific impulse, power efficiency, and life time. In this work, development of an anode layer Hall thruster (TAL) with magnetic field tailoring has been attempted. The TAL is possible to keep discharge in 1 to 2 kilovolts of anode voltage, which is useful to obtain high specific impulse. The magnetic field tailoring is used to minimize undesirable heat dissipation and secondary electron emission from the wall surrounding the plasma. We will report 3 W and 200 W thrusters performances measured by a pendulum thrust stand according to the magnetic field configuration. Also, the measured result will be compared with the plasma diagnostics conducted by an angular Faraday probe, a retarding potential analyzer, and a ExB probe.

  13. 20-mN Variable Specific Impulse (Isp) Colloid Thruster

    Science.gov (United States)

    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.

  14. Evaluation of externally heated pulsed MPD thruster cathodes

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    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.

  16. Power Electronics Development for the SPT-100 Thruster

    Science.gov (United States)

    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.

  17. NEXT Ion Thruster Performance Dispersion Analyses

    Science.gov (United States)

    Soulas, George C.; Patterson, Michael J.

    2008-01-01

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

  18. Description of the Prometheus Program Alternator/Thruster Integration Laboratory (ATIL)

    Science.gov (United States)

    Baez, Anastacio N.; Birchenough, Arthur G.; Lebron-Velilla, Ramon C.; Gonzalez, Marcelo C.

    2005-01-01

    The Project Prometheus Alternator Electric Thruster Integration Laboratory's (ATIL) primary two objectives are to obtain test data to influence the power conversion and electric propulsion systems design, and to assist in developing the primary power quality specifications prior to system Preliminary Design Review (PDR). ATIL is being developed in stages or configurations of increasing fidelity and complexity in order to support the various phases of the Prometheus program. ATIL provides a timely insight of the electrical interactions between a representative Permanent Magnet Generator, its associated control schemes, realistic electric system loads, and an operating electric propulsion thruster. The ATIL main elements are an electrically driven 100 kWe Alternator Test Unit (ATU), an alternator controller using parasitic loads, and a thruster Power Processing Unit (PPU) breadboard. This paper describes the ATIL components, its development approach, preliminary integration test results, and current status.

  19. Shared Magnetics Hall Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In the proposed Phase II program, Busek Co. will demonstrate an innovative methodology for clustering Hall thrusters into a high performance, very high power...

  20. Oxygen-Methane Thruster Project

    Data.gov (United States)

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

  1. Shared Magnetics Hall Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In the proposed Phase I program, Busek Co. will demonstrate an innovative methodology for clustering Hall thrusters into a high performance, very high power...

  2. Experimental results of an iodine plasma in PEGASES gridded thruster

    Science.gov (United States)

    Grondein, Pascaline; Aanesland, Ane

    2015-09-01

    In the electric gridded thruster PEGASES, both positive and negative ions are expelled after extraction from an ion-ion plasma. This ion-ion plasma is formed downstream a localized magnetic field placed a few centimeters from the ionization region, trapping and cooling down the electron to allow a better attachment to an electronegative gas. For this thruster concept, iodine has emerged as the most attractive option. Heavy, under diatomic form and therefore good for high thrust, its low ionization threshold and high electronegativity lead to high ion-ion densities and low RF power. After the proof-of-concept of PEGASES using SF6 as propellant, we present here experimental results of an iodine plasma studied inside PEGASES thruster. At solid state at standard temperature and pressure, iodine is heated to sublimate, then injected inside the chamber where the neutral gas is heated and ionized. The whole injection system is heated to avoid deposition on surfaces and a mass flow controller allows a fine control on the neutral gas mass flow. A 3D translation stage inside the vacuum chamber allows volumetric plasma studies using electrostatic probes. The results are also compared with the global model dedicated to iodine as propellant for electric gridded thrusters. This work has been done within the LABEX Plas@par project, and received financial state aid managed by the Agence Nationale de la Recherche, as part of the programme ``Investissements d'avenir.''

  3. Electric propulsion - A high energy capability for solar system exploration

    Science.gov (United States)

    Atkins, K. L.

    1976-01-01

    The principles of spacecraft electric (ion thruster) propulsion are briefly reviewed. Attention is given to the inner and outer planet applications of electric (and solar electric) propulsion. Electric propulsion is considered as a stepping stone to nuclear electric propulsion.

  4. MHD seawater thruster performance: A comparison of predictions with experimental results from a two Tesla test facility

    Energy Technology Data Exchange (ETDEWEB)

    Picologlou, B.F.; Doss, E.D.; Geyer, H.K. (Argonne National Lab., IL (United States)); Sikes, W.C.; Ranellone, R.F. (Newport News Shipbuilding and Dry Dock Co., VA (United States))

    1992-01-01

    A two Tesla test facility was designed, built, and operated to investigate the performance of magnetohydrodynamic (MHD) seawater thrusters. The results of this investigation are used to validate a design oriented MHD thruster performance computer code. The thruster performance code consists of a one-dimensional MHD hydrodynamic model coupled to a two-dimensional electrical model. The code includes major loss mechanisms affecting the performance of the thruster. Among these losses are the joule dissipation losses, frictional losses, electrical end losses, and single electrode potential losses. The facility test loop, its components, and their design are presented in detail. Additionally, the test matrix and its rationale are discussed. Representative experimental results of the test program are presented, and are compared to pretest computer model predictions. Good agreement between predicted and measured data has served to validate the thruster performance computer models.

  5. Carbon Back Sputter Modeling for Hall Thruster Testing

    Science.gov (United States)

    Gilland, James H.; Williams, George J.; Burt, Jonathan M.; Yim, John Tamin

    2016-01-01

    Lifetime requirements for electric propulsion devices, including Hall Effect thrusters, are continually increasing, driven in part by NASA's inclusion of this technology in it's exploration architecture. NASA will demonstrate high-power electric propulsion system on the Solar Electric Propulsion Technology Demonstration Mission (SEP TDM). The Asteroid Redirect Robotic mission is one candidate SEP TDM, which is projected to require tens of thousands of thruster life. As thruster life is increased, for example through the use of improved magnetic field designs, the relative influence of facility effects increases. One such effect is the sputtering and redeposition, or back sputter, of facility materials by the high energy thruster plumes. In support of wear testing for the Hall Effect Rocket with Magnetic Shielding (HERMeS) project, the back sputter from a Hall effect thruster plume has been modeled for the NASA Glenn Research Center's Vacuum Facility 5. The predicted wear at a near-worst case condition of 600 V, 12.5 kW was found to be on the order of 1 micron/kh in a fully carbon-lined chamber. A more detailed numerical Monte Carlo code was also modified to estimate back sputter for a detailed facility and pumping configuration. This code demonstrated similar back sputter rate distributions, but is not yet accurately modeling the magnitudes. The modeling has been benchmarked to recent HERMeS wear testing, using multiple microbalance measurements. These recent measurements have yielded values on the order of 1.5 - 2 micron/kh at 600 V and 12.5 kW.

  6. Preheating Cold Gas Thruster Flow Through a Thermal Energy Storage Conversion System

    Science.gov (United States)

    2013-01-01

    Journal Article 3. DATES COVERED (From - To) January 2013- October 2013 4. TITLE AND SUBTITLE Preheating Cold Gas Thruster Flow Through a Thermal Energy... Gas Thruster Flow through a Thermal Energy Storage Conversion System Michael R. Reid1 United States Air Force, Colorado Springs, CO, 80840 David B...specific impulse relative to a cold gas flow. Electric propulsion systems, the primary competitor to solar thermal propulsion systems, rely on the rather

  7. Iodine Hall Thruster for Space Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In the Phase I program, Busek Co. Inc. tested an existing Hall thruster, the BHT-8000, on iodine propellant. The thruster was fed by a high flow iodine feed system,...

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

    Science.gov (United States)

    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.

  9. Elementary scaling laws for the design of low and high power hall effect thrusters

    Science.gov (United States)

    Dannenmayer, K.; Mazouffre, S.

    2011-10-01

    An advanced set of scaling laws for Hall effect thrusters running with Xenon as propellant is established on the basis of the existence of an optimum atom number density that warrants a high efficiency thruster operation. A set of general relationships between macroscopic quantities, like thrust and input power, dimensions, including the channel length, the channel width and the channel mean diameter, and magnetic field strength are inferred from the main physical processes at work in a Hall thruster discharge. The "atom density constraint" of which the nature is here critically interpreted allows simplifying those relationships as it leads to a linear dependency between the channel length and mean diameter. Scaling laws which represent an essential tool for sizing up and down Hall thrusters are eventually obtained after proportionnality coefficients are determined. This last step is realized by means of a vast database that presently encompasses 33 single-stage Hall thrusters. In order to illustrate the usefulness of this new set of scaling laws, two practical applications are given and discussed. The scaling laws are first employed to calculate the dimensions and the operating parameters for a 20-kilowatt Hall thruster capable of producing 1 N of thrust. Such an electrical engine would permit orbit transfer of large communication satellites. Finally, the geometry of a Hall thruster is determined for tolerating 100 kW, an interesting power level for interplanetary trips.

  10. Mission and System Advantages of Iodine Hall Thrusters

    Science.gov (United States)

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

    2014-01-01

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

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

    Science.gov (United States)

    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.

  12. Helicon plasma thruster discharge model

    Energy Technology Data Exchange (ETDEWEB)

    Lafleur, T., E-mail: trevor.lafleur@lpp.polytechnique.fr [Laboratoire de Physique des Plasmas, CNRS, Sorbonne Universités, UPMC Univ Paris 06, Univ Paris-Sud, Ecole Polytechnique, 91128 Palaiseau, France and ONERA - The French Aerospace Lab, 91120 Palaiseau (France)

    2014-04-15

    By considering particle, momentum, and energy balance equations, we develop a semi-empirical quasi one-dimensional analytical discharge model of radio-frequency and helicon plasma thrusters. The model, which includes both the upstream plasma source region as well as the downstream diverging magnetic nozzle region, is compared with experimental measurements and confirms current performance levels. Analysis of the discharge model identifies plasma power losses on the radial and back wall of the thruster as the major performance reduction factors. These losses serve as sinks for the input power which do not contribute to the thrust, and which reduce the maximum plasma density and hence propellant utilization. With significant radial plasma losses eliminated, the discharge model (with argon) predicts specific impulses in excess of 3000 s, propellant utilizations above 90%, and thruster efficiencies of about 30%.

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

    Science.gov (United States)

    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.

  14. Cathode-less gridded ion thrusters for small satellites

    Science.gov (United States)

    Aanesland, Ane

    2016-10-01

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

  15. Modelling and Simulation of Variable Speed Thruster Drives with Full-Scale Verification

    Directory of Open Access Journals (Sweden)

    Jan F. Hansen

    2001-10-01

    Full Text Available In this paper considerations about modelling and simulation of variable speed thruster drives are made with comparison to full scale measurements from Varg FPSO. For special purpose vessels with electric propulsion operating in DP (Dynamic Positioning mode the thruster drives are essential for the vessel operation. Different model strategies of thruster drives are discussed. An advanced thruster drive model with a dynamic motor model and field vector control principle is shown. Simulations are performed with both the advanced model and a simplified model. These are compared with full-scale measurements from Varg FPSO. The simulation results correspond well with the measurements, for both the simplified model and the advanced model.

  16. On the microscopic mechanism of ion-extraction of a gridded ion propulsion thruster

    CERN Document Server

    Kirmse, Danny

    2013-01-01

    The following paper includes a physical microscopic particle-description of the phenomena and mechanisms that lead to the extraction of ions with the aim to generate thrust. This theoretical treatise arose from the intention to visualize the behavior of the involved particles under effect of the involved electrical fields. By this way, an underlying basis for experimental investigations of the work of an ion thruster should be formed. So a foundation was created, which explains the ion extracting and so thrust generating function of an ion thruster. The theoretical work was related to the Radio-frequency Ion Thruster (RIT). But the model worked out can be generalized for all thruster types that use electrostatic fields to extract positively charged ions.

  17. Electric propulsion and its applications to space missions

    Science.gov (United States)

    Finke, R. C.

    1981-01-01

    Consideration is given the NASA, Air Force and European electric propulsion programs, the characteristics of primary electric propulsion systems, nuclear electric orbit transfer vehicles, and such topics in the fundamental processes of electrostatic thrusters as sputtering in mercury ion thrusters, the screen hole plasma sheath of an ion accelerator system, and the modelling of ion beam neutralization and nitrogen chemisorption. Also considered are electrostatic thruster components and systems, electromagnetic thrusters such as MPD and RIT systems, electric rail guns and mass drivers, power sources which include solar and nuclear alternatives, power conversion systems and their cooling apparatus, and the environmental interactions between spacecraft and their electric propulsion systems.

  18. Electric Propulsion

    Science.gov (United States)

    Baggett, R.

    2004-11-01

    Next Generation Electric Propulsion (NGEP) technology development tasks are working towards advancing solar-powered electric propulsion systems and components to levels ready for transition to flight systems. Current tasks within NGEP include NASA's Evolutionary Xenon Thruster (NEXT), Carbon Based Ion Optics (CBIO), NSTAR Extended Life Test (ELT) and low-power Hall Effect thrusters. The growing number of solar electric propulsion options provides reduced cost and flexibility to capture a wide range of Solar System exploration missions. Benefits of electric propulsion systems over state-of-the-art chemical systems include increased launch windows, which reduce mission risk; increased deliverable payload mass for more science; and a reduction in launch vehicle size-- all of which increase the opportunities for New Frontiers and Discovery class missions. The Dawn Discovery mission makes use of electric propulsion for sequential rendezvous with two large asteroids (Vesta then Ceres), something not possible using chemical propulsion. NEXT components and thruster system under development have NSTAR heritage with significant increases in maximum power and Isp along with deep throttling capability to accommodate changes in input power over the mission trajectory. NEXT will produce engineering model system components that will be validated (through qualification-level and integrated system testing) and ready for transition to flight system development. NEXT offers Discovery, New Frontiers, Mars Exploration and outer-planet missions a larger deliverable payload mass and a smaller launch vehicle size. CBIO addresses the need to further extend ion thruster lifetime by using low erosion carbon-based materials. Testing of 30-cm Carbon-Carbon and Pyrolytic graphite grids using a lab model NSTAR thruster are complete. In addition, JPL completed a 1000 hr. life test on 30-cm Carbon-Carbon grids. The NSTAR ELT was a life time qualification test started in 1999 with a goal of 88 kg

  19. J series thruster thermal test results

    Science.gov (United States)

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

    1982-01-01

    Test experience with J series ion thrusters have indicated that the present thruster design may result in excessive temperatures in areas which utilize organic materials such as wire insulation, with the resultant outgassing and potential contamination of insulating materials. Further, it appears that thermal data obtained with earlier thruster designs, such as the 700 series thruster, may not be directly applicable to the J series design. Two J series thrusters were fitted with thermocouples and critical temperatures measured for a variety of configurations and operating parameters. Completely enclosing the thruster to reduce facility contamination significantly increased temperatures prompting the selection of a compromise geometry for life testing. The operating parameter having the largest effect on temperatures was discharge power, while beam power affected little else than extraction system temperatures. Several off-normal operating modes were also investigated. Data believed to be sufficient to effectively modify existing thermal models were obtained from the tests.

  20. Electron-wall Interaction in Hall Thrusters

    Energy Technology Data Exchange (ETDEWEB)

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

    2005-02-11

    Electron-wall interaction effects in Hall thrusters are studied through measurements of the plasma response to variations of the thruster channel width and the discharge voltage. The discharge voltage threshold is shown to separate two thruster regimes. Below this threshold, the electron energy gain is constant in the acceleration region and therefore, secondary electron emission (SEE) from the channel walls is insufficient to enhance electron energy losses at the channel walls. Above this voltage threshold, the maximum electron temperature saturates.

  1. Low-Voltage Hall Thruster Mode Transitions

    Science.gov (United States)

    2014-06-01

    Technical Paper 3. DATES COVERED (From - To) June 2014- July 2014 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER In-House Low-Voltage Hall Thruster Mode...ABSTRACT Past investigations of the 6kW-class H6 Hall thruster during low-voltage operation revealed two operating modes, corresponding to the...topologies were characterized for the H6 Hall thruster from 100V to 200V discharge, with variation in cathode flow fraction, cathode position inside and

  2. Advanced Microwave Electrothermal Thruster (AMET) Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Orbital Technologies Corporation (ORBITEC) and the University of Alabama at Huntsville (UAH) propose to develop the Advanced Microwave Electrothermal Thruster...

  3. NASA Brief: Q-Thruster Physics

    Science.gov (United States)

    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.

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

    Science.gov (United States)

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

  5. Coaxial microwave electrothermal thruster performance in hydrogen

    Science.gov (United States)

    Richardson, W.; Asmussen, J.; Hawley, M.

    1994-01-01

    The microwave electro thermal thruster (MET) is an electric propulsion concept that offers the promise of high performance combined with a long lifetime. A unique feature of this electric propulsion concept is its ability to create a microwave plasma discharge separated or floating away from any electrodes or enclosing walls. This allows propellant temperatures that are higher than those in resistojets and reduces electrode and wall erosion. It has been demonstrated that microwave energy is coupled into discharges very efficiently at high input power levels. As a result of these advantages, the MET concept has been identified as a future high power electric propulsion possibility. Recently, two additional improvements have been made to the coaxial MET. The first was concerned with improving the microwave matching. Previous experiments were conducted with 10-30 percent reflected power when incident power was in excess of 600 W(exp 6). Power was reflected back to the generator because the impedance of the MET did not match the 50 ohm impedance of the microwave circuit. To solve this problem, a double stub tuning system has been inserted between the MET and the microwave power supply. The addition of the double stub tuners reduces the reflected power below 1 percent. The other improvement has prepared the coaxial MET for hydrogen experiments. To operate with hydrogen, the vacuum window which separates the coaxial line from the discharge chamber has been changed from teflon to boron nitride. All the microwave energy delivered to the plasma discharge passes through this vacuum window. This material change had caused problems in the past because of the increased microwave reflection coefficients associated with the electrical properties of boron nitride. However, by making the boron nitride window electrically one-half of a wavelength long, power reflection in the window has been eliminated. This technical note summarizes the experimental performance of the improved

  6. First Firing of a 100-kW Nested-Channel Hall Thruster

    Science.gov (United States)

    2013-09-01

    need to complete their mission (i.e., SEP cargo vessel ) drive the requirement for high-power (>100 kW) EP. N The 33st International Electric ...Frisbee, R., ““Evaluation of High-power Solar Electric Propulsion Using Advanced Ion, Hall, MPD, and PIT Thrusters for Lunar and Mars Cargo ...Approved for Public Release; Distribution Unlimited. PA#13553 13. SUPPLEMENTARY NOTES Conference paper for the 33rd International Electric Propulsion

  7. Performance characterization of a permanent-magnet helicon plasma thruster

    Science.gov (United States)

    Takahashi, Kazunori; Charles, Christine; Boswell, Rod

    2012-10-01

    Helicon plasma thrusters operated at a few kWs of rf power is an active area of an international research. Recent experiments have clarified part of the thrust-generation mechanisms. Thrust components which have been identified include an electron pressure inside the source region and a Lorentz force due to an electron diamagnetic drift current and a radial component of the applied magnetic field. The use of permanent magnets (PMs) instead of solenoids is one of the solutions for improving the thruster efficiency because it does not require electricity for the magnetic nozzle formation. Here the thrust imparted from a permanent-magnet helicon plasma thruster is directly measured using a pendulum thrust balance. The source consists of permanent magnet (PM) arrays, a double turn rf loop antenna powered by a 13.56 MHz rf generator and a glass source tube. The PM arrays provide a magnetic nozzle near the open exit of the source and two configurations, which have maximum field strengths of about 100 and 270 G, are tested. A thrust of 15 mN, specific impulse of 2000 sec and a thrust efficiency of 8 percent are presently obtained for 2 kW of input power, 24 sccm flow rate of argon and the stronger magnetic field configuration.

  8. High-Efficiency Hall Thruster Discharge Power Converter

    Science.gov (United States)

    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.

  9. Development of ion thruster IT-500

    Science.gov (United States)

    Koroteev, Anatoly S.; Lovtsov, Alexander S.; Muravlev, Vyacheslav A.; Selivanov, Mikhail Y.; Shagayda, Andrey A.

    2017-05-01

    A high-power ion thruster IT-500 was designed, manufactured and tested at Keldysh Research Center within a transport-power module project. This module is being designed to perform near-Earth space and interplanetary transport missions. In its nominal operation mode, IT-500 provides thrust in the range from 375 to 750 mN at specific impulse of 70 000 m/s and thrust efficiency of 0.75. Due to a high cost of the experimental testing of a large thruster, the emphasis was placed on the numerical optimization of the thruster design. The thruster completed performance tests and a 300 h wear test. The output characteristics of the thruster, obtained during the tests, confirmed the correctness of the provisional numerical optimization. IT-500 design, performance, and validation of the design approaches are discussed in this paper. Contribution to the Topical Issue: "Physics of Ion Beam Sources", edited by Holger Kersten and Horst Neumann.

  10. Coil system for plasmoid thruster

    Science.gov (United States)

    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.

  11. BMDO electric space-propulsion program

    Science.gov (United States)

    Caveny, Leonard H.; Curran, Francis M.; Brophy, John R.

    1993-01-01

    Electric propulsion (EP) applications being considered include: orbit insertion, orbit repositioning, station keeping, and elusive maneuvering. Typically, 1 to 5 kW are available for EP. A class of thrusters, the Hall-effect thrusters, is extensively researched, developed and flown by Russia. These thrusters, using xenon propellant, perform reliably, e.g., at 1.35 kW, 600 s specific impulse, 50 percent efficiency and greater than 2000 hr life. This specific impulse and efficiency combination is superior to the present arcjets for several Ballistic Missile Defense Organization (BMDO) applications. Three versions of the Hall thrusters are part of the experimental evaluation. Since today's goals are within reach of available thrusters and power sources, emphasis is being placed on such topics as: thruster lifetime, spacecraft interactions, electromagnetic interference, and erosion product deposition. Facilities in U.S. laboratories are being specially configured to achieve these goals.

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

    Science.gov (United States)

    Rawlin, V. K.

    1979-01-01

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

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

    Science.gov (United States)

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

    2015-11-01

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

  14. Non-Contact Thermal Characterization of NASA's HERMeS Hall Thruster

    Science.gov (United States)

    Huang, Wensheng; Kamhawi, Hani; Meyers, James L.; Yim, John T.; Neff, Gregory

    2015-01-01

    The Thermal Characterization Test of NASAs 12.5-kW Hall thruster is being completed. This thruster is being developed to support of a number of potential Solar Electric Propulsion Technology Demonstration Mission concepts, including the Asteroid Redirect Robotic Mission concept. As a part of this test, an infrared-based, non-contact thermal imaging system was developed to measure Hall thruster surfaces that are exposed to high voltage or harsh environment. To increase the accuracy of the measurement, a calibration array was implemented, and a pilot test was performed to determine key design parameters for the calibration array. The raw data is analyzed in conjunction with a simplified thermal model of the channel to account for reflection. The reduced data will be used to refine the thruster thermal model, which is critical to the verification of the thruster thermal specifications. The present paper will give an overview of the decision process that led to identification of the need for a non-contact temperature diagnostic, the development of said diagnostic, the measurement results, and the simplified thermal model of the channel.

  15. Human Outer Solar System Exploration via Q-Thruster Technology

    Science.gov (United States)

    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

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

    Science.gov (United States)

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

    2012-01-01

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

  17. Iodine Hall Thruster Propellant Feed System for a CubeSat

    Science.gov (United States)

    Polzin, Kurt A.

    2014-01-01

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

  18. Light Metal Propellant Hall Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop light metal Hall Effect thrusters that will help reduce the travel time, mass, and cost of SMD spacecraft. Busek has identified three...

  19. Precision Electrospray Thruster Assembly (PETA) Project

    Data.gov (United States)

    National Aeronautics and Space Administration — New low cost, low volume, low power, rugged electrospray thrusters will be ideal as actuators for precision thrusting, if provided with precision high voltage power...

  20. T6 Ion Thruster Technology Development Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Provide discharge chamber and grid modeling for the new T6 based on JPL expertise on ion thruster performance and life; Enable/guide the T6 upgrade development to...

  1. Dual Mode Low Power Hall Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Sample and return missions desire and missions like Saturn Observer require a low power Hall thruster that can operate at high thrust to power as well as high...

  2. High Thrust Efficiency MPD Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Magnetoplasmadynamic (MPD) thrusters can provide the high-specific impulse, high-power propulsion required to support human and robotic exploration missions to the...

  3. Optimized Magnetic Nozzles for MPD Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Magnetoplasmadynamic (MPD) thrusters can provide the high-specific impulse, high-power propulsion required to enable ambitious human and robotic exploration missions...

  4. Advanced High Efficiency Durable DACS Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Systima is developing a high performance 25 lbf DACS thruster that operates with a novel non-toxic monopropellant. The monopropellant has a 30% higher...

  5. Acoustic Resonance Reaction Control Thruster (ARCTIC) Project

    Data.gov (United States)

    National Aeronautics and Space Administration — ORBITEC proposes to develop and demonstrate the innovative Acoustic Resonance Reaction Control Thruster (ARCTIC) to provide rapid and reliable in-space impulse...

  6. Multiscale Modeling of Hall Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — New multiscale modeling capability for analyzing advanced Hall thrusters is proposed. This technology offers NASA the ability to reduce development effort of new...

  7. MPD thruster research issues, activities, strategies

    Science.gov (United States)

    1991-01-01

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

  8. Colloid Thrusters, Physics, Fabrication and Performance

    Science.gov (United States)

    2005-11-17

    response, including the time for reviewing in. tata needed, and completing and reviewing this collection of information. Send comments regarding this...a discussion with colleagues during the 2nd Colloid Thruster/ Nano Electrojet Workshop (MIT, April 14- 15, 2005, Ref. [11]) an agreement was reached...23 Jul 2003. 11. Second Colloid Thruster/ Nano Electrojet Workshop, CD with a collection of presentations by attendees to this Workshop. MIT, April 14

  9. MEMS-Based Solid Propellant Rocket Array Thruster

    Science.gov (United States)

    Tanaka, Shuji; Hosokawa, Ryuichiro; Tokudome, Shin-Ichiro; Hori, Keiichi; Saito, Hirobumi; Watanabe, Masashi; Esashi, Masayoshi

    The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3 4 W with RK and 4 6 W without RK. Measured impulse thrusts were from 2×10-5 Ns to 3×10-4 Ns after the calculation of compensation for air dumping.

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

    Science.gov (United States)

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

    2015-01-01

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

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

    Science.gov (United States)

    Pencil, Eric J.; Benson, Scott W.

    2008-01-01

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

  12. Implementation and Initial Validation of a 100-Kilowatt Class Nested-Channel Hall Thruster

    Science.gov (United States)

    Hall, Scott J.; Florenz, Roland E.; Gallimore, Alec D.; Kamhawi, Hani; Brown, Daniel L.; Polk, James E.; Goebel, Dan; Hofer, Richard R.

    2014-01-01

    The X3 is a 100-kilowatt class nested-channel Hall thruster developed by the Plasmadynamics and Electric Propulsion Laboratory at the University of Michigan in collaboration with the Air Force Research Laboratory and NASA. The cathode, magnetic circuit, boron nitride channel rings, and anodes all required specific design considerations during thruster development, and thermal modeling was used to properly account for thermal growth in material selection and component design. A number of facility upgrades were required at the University of Michigan to facilitate operation of the X3. These upgrades included a re-worked propellant feed system, a completely redesigned power and telemetry break-out box, and numerous updates to thruster handling equipment. The X3 was tested on xenon propellant at two current densities, 37% and 73% of the nominal design value. It was operated to a maximum steady-state discharge power of 60.8 kilowatts. The tests presented here served as an initial validation of thruster operation. Thruster behavior was monitored with telemetry, photography and high-speed current probes. The photography showed a uniform plume throughout testing. At constant current density, reductions in mass flow rate of 18% and 26% were observed in the three-channel operating configuration as compared to the superposition of each channel running individually. The high-speed current probes showed that the thruster was stable at all operating points and that the channels influence each other when more than one is operating simultaneously. Additionally, the ratio of peak-to-peak AC-coupled discharge current oscillations to mean discharge current did not exceed 51% for any operating points reported here, and did not exceed 17% at the higher current density.

  13. Control capability analysis for complex spacecraft thruster configurations

    Institute of Scientific and Technical Information of China (English)

    2010-01-01

    The set of forces and moments that can be generated by thrusters of a spacecraft is called the"control capability"with respect to the thruster configuration.If the control capability of a thruster configuration is adequate to fulfill a given space mission,we say this configuration is a feasible one with respect to the task.This study proposed a new way to analyze the control capability of the complex thruster configuration.Precise mathematical definitions of feasibility were proposed,based on which a criterion to judge the feasibility of the thruster configuration was presented through calculating the shortest distance to the boundary of the controllable region as a function of the thruster configuration.Finally,control capability analysis for the complex thruster configuration based on its feasibility with respect to the space mission was given followed by a 2-D thruster configuration example to demonstrate its validity.

  14. Performance Characterization of a Three-Axis Hall Effect Thruster

    Science.gov (United States)

    2010-12-01

    here represents the first efforts to operate and quantify the performance of a three-axis Hall effect thruster. This thruster is based on the Busek BHT ...thruster were developed and thrust and current density measurements were performed and compared with the baseline BHT -200. The three-axis thruster was...efficiencies than the BHT -200. Beam current density measurements conducted using a guarded Faraday probe showed significant differences in plume divergence

  15. Design and operations of Hall thruster with segmented electrodes

    Energy Technology Data Exchange (ETDEWEB)

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

    1999-12-10

    Principles of the Hall thruster with segmented electrodes are explored. A suitable vacuum facility was put into service. For purposes of comparison between segmented and conventional thruster approaches, a modular laboratory prototype thruster was designed and built. Under conventional operation, the thruster achieves state-of-the-art efficiencies (56% at 300 V and 890 W). Very preliminary results under operation with segmented electrodes are also described.

  16. Design and Operation of Hall Thruster with Segmented Electrodes

    Energy Technology Data Exchange (ETDEWEB)

    A.A. Litvak; L.A. Dorf; N.J. Fisch; Y. Raitses

    1999-07-01

    Principles of the Hall thruster with segmented electrodes are explored. A suitable vacuum facility was put into service. For purposes of comparison between segmented and conventional thruster approaches, a modular laboratory prototype thruster was designed and built. Under conventional operation, the thruster achieves state-of-the-art efficiencies (56% at 300 V and 890 W). Very preliminary results under operation with segmented electrodes are also described.

  17. Investigations of Probe Induced Perturbations in a Hall Thruster

    Energy Technology Data Exchange (ETDEWEB)

    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.

  18. Coaxial plasma thrusters for high specific impulse propulsion

    Science.gov (United States)

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

    1991-01-01

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

  19. Parametric Investigations of Non-Conventional Hall Thruster

    Energy Technology Data Exchange (ETDEWEB)

    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.

  20. OL-AC Phillips Laboratory MPD thruster research program

    Science.gov (United States)

    Tilley, Dennis L.

    1992-01-01

    The topics are presented in viewgraph form and include the following: facility construction; quadruple langmuir probe measurements; hollow/porous anode magnetoplasmadynamic (MPD) thruster; the measurement of the ionization fraction inside of the MPD thruster; and the experimental investigation of the effects of microturbulence on MPD thruster performance.

  1. Spatial characterization of electric potentials generated by pulsed microelectrode arrays.

    Science.gov (United States)

    Kandagor, V; Cela, C J; Sanders, C A; Greenbaum, E; Lazzi, G; Humayun, M S; Zhou, D M; Castro, R; Gaikwad, S; Little, J

    2010-01-01

    This presentation is a report on the in situ characterization of stimulating microelectrodes in the context of multielectrode retinal prosthetic implants. The experimental system approximately replicates the geometric and electrical parameters of Second Sight Medical Products' Argus II Retinal Implant. Topographic maps of electric potentials have been prepared for a 60 electrode structure in which selected electrodes were stimulated with biphasic repetitively pulsed charge densities at 100 microC·cm(-2). Surface contour maps were prepared using a 10 microm diameter recording electrode.

  2. Electrostatic ion thrusters - towards predictive modeling

    Energy Technology Data Exchange (ETDEWEB)

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

    2014-02-15

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

  3. Enhancing Micro-Cathode Arc Thruster (muCAT) Plasma Generation to Analyze Magnetic Field Angle Effects on Sheath Formation in Hall Thrusters

    Science.gov (United States)

    Lukas, Joseph Nicholas

    Using a Delta IV or Atlas V launch vehicle to send a payload into Low Earth Orbit can cost between 13,000 and 14,000 per kilogram. With payloads that utilize a propulsion system, maximizing the efficiency of that propulsion system would not only be financially beneficial, but could also increase the range of possible missions and allow for a longer mission lifetime. This dissertation looks into efficiency increases in the Micro-Cathode Arc Thruster (muCAT) and Hall Thruster. The muCAT is an electric propulsion device that ablates solid cathode material, through an electrical arc discharge, to create plasma and ultimately produce thrust. About 90% of the arc discharge current is conducted by electrons, which go toward heating the anode and contribute very little to thrust, with only the remaining 10% going toward thrust in the form of ion current. I will discuss the results of an experiment in which electron heating on a low melting point anode was shown to increase ion current, which theoretically should increase thrust levels at low frequencies. Another feature of the muCAT is the use of an external magnetic solenoid which increases thrust, ion current, and causes uniform cathode erosion. An experiment has shown that efficiency can also be increased by removing the external magnetic field power supply and, instead, utilizing the residual arc current to power the magnetic solenoid. A Hall Thruster is a type of electric propulsion device that accelerates ions across an electric potential between an anode and magnetically trapped electrons. The limiting factor in Hall Thruster operation is the lifetime of the wall material. During operation, a positively charged layer forms over the surface of the walls, known as a plasma sheath, which contributes to wall erosion. Therefore, by reducing or eliminating the sheath layer, Hall Thruster operational lifetime can increase. Computational modeling has shown that large magnetic field angles and large perpendicular electric

  4. A brief review of JPL's electric propulsion technology activities

    Science.gov (United States)

    Barnett, John W.; Chopra, Ann; Deininger, William D.; Garner, Charles E.; Pivirotto, Thomas J.; Sercel, Joel C.

    1989-01-01

    Near-term objectives and recent technological progress of JPL's electric propulsion program are discussed. Particular attention is given to accomplishments for ion, magnetoplasmadynamic (MPD), electron-cyclotron resonance (ECR), and arcjet thrusters. Xenon ion thruster erosion tests indicate a 15-fold reduction in tantalum baffle erosion when nitrogen is added to the xenon propellant and steady-state cylindrical MPD thruster tests at powers up to 72 kW show distinct self-constricted and diffuse discharge modes. An ECR thruster was operated at up to 7 kW with plasma acceleration at energies up to 7 kW; there was plasma acceleration at energies approaching 100 electron volts.

  5. Experimental Study of the Microdischarge Plasma Thruster (MDPT)

    Science.gov (United States)

    Kc, Utsav; Varghese, Philip; Raja, Laxminarayan

    2008-10-01

    Small satellite propulsion requirements dictate the need for a scaled down propulsion device capable of providing low thrust with small impulse bits. We have designed and studied a simple miniaturized thruster called Microdischarge Plasma Thruster (MDPT). It comprises a tri-layer sandwich structure with a dielectric layer sandwiched between two electrode layers, and a contoured through hollow drilled into the structure. Each layer is 100's microns in thickness and the hole diameter of the same order. Argon is used as the propellant gas with flow rates of ˜ 1 SCCM. The pressure is adequate to produce a stable microdischarge between the electrodes even with modest voltages (˜1000 V). The microdischarge adds heat to the supersonic portion of the flowing gas which is shown to produce additional thrust over the baseline cold gas flow. The studies have also demonstrated that the MDPT exhaust plume is composed of ions albeit at low concentrations, suggesting possibility of MDPT to be operated in a mixed electrothermal/electrostatic mode. We present discussion of multiple discharge operating modes and electrical characteristics of the MDPT. Spectral measurements of the plume are used to determine its composition and calculate its temperature. The momentum thrust of the MDPT is measured with a torsion balance.

  6. Optimization of energy transfer in microwave electrothermal thrusters

    Science.gov (United States)

    Sullivan, D. J.; Micci, M. M.

    1993-01-01

    Results are presented from preliminary tests conducted to evaluate the performance of a prototype microwave electrothermal thruster. The primary component of the device is a microwave resonant cavity. The device produces stable axial plasmas within a pressurized section of the cavity with the plasma positioned in the inlet region of the nozzle. Plasma stability is enhanced by axial power coupling, an optimal distribution of electric power density within the cavity, and a propellant gas flow which has a large vortical velocity component. The thruster has been operated with a number of propellant gases: helium, nitrogen, ammonia, and hydrogen. Plasmas can be formed in a reliable manner at cavity pressures of 1 kPa and incident power levels ranging from 50 W to 350 W, depending on the gas used, and can be operated at pressures up to 300 kPa at power levels up to 2200 W. Ideal performance results of vacuum Isp and thermal efficiency vs. specific power are presented for each gas. Representative results of this preliminary work are: He - Isp = 625 s, eta-thermal = 90 percent; N2 - Isp = 270 s, eta-thermal = 41 percent; NH3 - Isp = 475 s, eta-thermal= 55 percent; H2 - Isp = 1040 s, eta-thermal = 53 percent.

  7. Facility Effect Characterization Test of NASA's HERMeS Hall Thruster

    Science.gov (United States)

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

    2016-01-01

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

  8. Rotating plasma structures in the cross-field discharge of Hall thrusters

    Science.gov (United States)

    Mazouffre, Stephane; Grimaud, Lou; Tsikata, Sedina; Matyash, Konstantin

    2016-09-01

    Rotating plasma structures, also termed rotating spokes, are observed in various types of low-pressure discharges with crossed electric and magnetic field configurations, such as Penning sources, magnetron discharges, negative ion sources and Hall thrusters. Such structures correspond to large-scale high-density plasma blocks that rotate in the E×B drift direction with a typical frequency on the order of a few kHz. Although such structures have been extensively studied in many communities, the mechanism at their origin and their role in electron transport across the magnetic field remain unknown. Here, we will present insights into the nature of spokes, gained from a combination of experiments and advanced particle-in-cell numerical simulations that aim at better understanding the physics and the impact of rotating plasma structures in the ExB discharge of the Hall thruster. As rotating spokes appear in the ionization region of such thrusters, and are therefore difficult to probe with diagnostics, experiments have been performed with a wall-less Hall thruster. In this configuration, the entire plasma discharge is pushed outside the dielectric cavity, through which the gas is injected, using the combination of specific magnetic field topology with appropriate anode geometry.

  9. Bi-directional thruster development and test report

    Science.gov (United States)

    Jacot, A. D.; Bushnell, G. S.; Anderson, T. M.

    1990-01-01

    The design, calibration and testing of a cold gas, bi-directional throttlable thruster are discussed. The thruster consists of an electro-pneumatic servovalve exhausting through opposite nozzles with a high gain pressure feedback loop to optimize performance. The thruster force was measured to determine hysteresis and linearity. Integral gain was used to maximize performance for linearity, hysteresis, and minimum thrust requirements. Proportional gain provided high dynamic response (bandwidth and phase lag). Thruster performance is very important since the thrusters are intended to be used for active control.

  10. Mode Transitions in Hall Effect Thrusters

    Science.gov (United States)

    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.

  11. Recent Advances in Nuclear Powered Electric Propulsion for Space Exploration

    Science.gov (United States)

    Cassady, R. Joseph; Frisbee, Robert H.; Gilland, James H.; Houts, Michael G.; LaPointe, Michael R.; Maresse-Reading, Colleen M.; Oleson, Steven R.; Polk, James E.; Russell, Derrek; Sengupta, Anita

    2007-01-01

    Nuclear and radioisotope powered electric thrusters are being developed as primary in-space propulsion systems for potential future robotic and piloted space missions. Possible applications for high power nuclear electric propulsion include orbit raising and maneuvering of large space platforms, lunar and Mars cargo transport, asteroid rendezvous and sample return, and robotic and piloted planetary missions, while lower power radioisotope electric propulsion could significantly enhance or enable some future robotic deep space science missions. This paper provides an overview of recent U.S. high power electric thruster research programs, describing the operating principles, challenges, and status of each technology. Mission analysis is presented that compares the benefits and performance of each thruster type for high priority NASA missions. The status of space nuclear power systems for high power electric propulsion is presented. The paper concludes with a discussion of power and thruster development strategies for future radioisotope electric propulsion systems,

  12. Performance and Qualification of the Power Supply and Control Unit for the HEMP Thruster

    Science.gov (United States)

    Brag, R.; Herty, F.

    2014-08-01

    In 2013, Astrium GmbH delivered several flight model electronics for Electric Propulsion (EP) systems or corresponding components. One of the elements is a Power Supply and Control Unit (PSCU) for the Thales development "High Efficiency Multistage Plasma Thruster" (HEMP-T) (see Figure 1). This paper presents the PSCU specification and results of the qualification and acceptance phase of the EQM and the PFM.

  13. Measurements of Plasma Potential Distribution in Segmented Electrode Hall Thruster

    Energy Technology Data Exchange (ETDEWEB)

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

    2001-10-16

    Use of a segmented electrode placed at the Hall thruster exit can substantially reduce the voltage potential drop in the fringing magnetic field outside the thruster channel. In this paper, we investigate the dependence of this effect on thruster operating conditions and segmented electrode configuration. A fast movable emissive probe is used to measure plasma potential in a 1 kW laboratory Hall thruster with semented electrodes made of a graphite material. Relatively small probe-induced perturbations of the thruster discharge in the vicinity of the thruster exit allow a reasonable comparison of the measured results for different thruster configurations. It is shown that the plasma potential distribution is almost not sensitive to changes of the electrode potential, but depends on the magnetic field distribution and the electrode placement.

  14. A Numerical Study on Hydrodynamic Interactions between Dynamic Positioning Thrusters

    Energy Technology Data Exchange (ETDEWEB)

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

    2017-06-15

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

  15. Pulsed Plasma Thruster plume analysis

    Energy Technology Data Exchange (ETDEWEB)

    Parker, K. [Washington Univ., Aerospace and Energetics Research Program, Seattle, WA (United States)

    2003-11-01

    Micro-Pulsed Plasma Thrusters ({mu}PPTs) are a promising method for precision attitude control for small spacecraft in formation flying. They create an ionized plasma plume, which may interfere with other spacecraft in the formation. To characterize the ions in the plume, a diagnostic has been built that couples a drift tube with an energy analyzer. The drift tube provides time of flight measurements to determine the exhaust velocity, and the energy analyzer discriminates the ion energies. The energy analyzer measures the current on a collector plate downstream of four grids that repel electrons and ions below a specified energy. The first grid lowers the density of the plasma, therefore increasing Debye length. The second and fourth grids have a negative potential applied to them so they repel the electrons, while the third grid's voltage can be varied to repel lower energy ions. The ion energies can be computed by differentiating the data. Combining the information of the ion energies and their velocities identifies the ion masses in the PPT plume. The PPT used for this diagnostic is the micro-PPT developed for the Dawgstar satellite. This PPT uses 5.2 Joules per pulse and has a 2.3 cm{sup 2} propellant area, a 1.3 cm electrode length, and an estimated thrust of 85 {mu}N [C. Rayburn et al., AIAA-2000-3256]. This paper will describe the development and design of the time of flight/gridded energy analyzer diagnostic and present recent experimental results. (Author)

  16. Modelling a stationary plasma thruster for satellites; Modelisation d'un propulseur a plasma stationnaire pour satellites

    Energy Technology Data Exchange (ETDEWEB)

    Garrigues, L.

    1998-07-01

    Stationary plasma thrusters (SPT) are small propulsion systems with interesting properties for low orbit changes and N-S and E-W corrections of satellites. The functioning principle is based on the creation of a plasma outside the stationary equilibrium and under a magnetic field perpendicular to the axis of the discharge which leads to the generation of a ion beam used to propel the satellite. The French Stentor satellite project will use SPT-type thrusters. The aim of this work is to better understand the physical phenomena occurring in SPTs using numerical models. A first step has been the elaboration of a Monte Carlo particle model for the analysis of electrons transport inside the thruster and threw a microscopic approach. In a second step, the electrical characteristics (low frequency oscillations of the discharge current, plasma evolution) and the thruster performances (thrust, specific pulse and efficiency) are analyzed. A 1-D, quasi-neutral, transient and self-consistent (fluid and hybrid approaches) model has been elaborated which allows to follow the evolution of the discharge in the channel. Thanks to the use of simplification hypotheses, complete and various studies about the influence of external parameters on the characteristics of the thruster could have been performed (flow rate of injected gas, potential applied, shape and value of the magnetic field). Results are qualitatively in agreement with the experimental results and with results obtained with other models. (J.S)

  17. Overview of NASA Iodine Hall Thruster Propulsion System Development

    Science.gov (United States)

    Smith, Timothy D.; Kamhawi, Hani; Hickman, Tyler; Haag, Thomas; Dankanich, John; Polzin, Kurt; Byrne, Lawrence; Szabo, James

    2016-01-01

    NASA is continuing to invest in advancing Hall thruster technologies for implementation in commercial and government missions. The most recent focus has been on increasing the power level for large-scale exploration applications. However, there has also been a similar push to examine applications of electric propulsion for small spacecraft in the range of 300 kg or less. There have been several recent iodine Hall propulsion system development activities performed by the team of the NASA Glenn Research Center, the NASA Marshall Space Flight Center, and Busek Co. Inc. In particular, the work focused on qualification of the Busek 200-W BHT-200-I and development of the 600-W BHT-600-I systems. This paper discusses the current status of iodine Hall propulsion system developments along with supporting technology development efforts.

  18. Micro Cathode Arc Thruster for PhoneSat: Development and Potential Applications

    Science.gov (United States)

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

    2014-01-01

    NASA Ames Research Center and the George Washington University are developing an electric propulsion subsystem that will be integrated into the PhoneSat bus. Experimental tests have shown a reliable performance by firing three different thrusters at various frequencies in vacuum conditions. The interface consists of a microcontroller that sends a trigger pulse to the Pulsed Plasma Unit that is responsible for the thruster operation. A Smartphone is utilized as the main user interface for the selection of commands that control the entire system. The propellant, which is the cathode itself, is a solid cylinder made of Titanium. This simplicity in the design avoids miniaturization and manufacturing problems. The characteristics of this thruster allow an array of µCATs to perform attitude control and orbital correction maneuvers that will open the door for the implementation of an extensive collection of new mission concepts and space applications for CubeSats. NASA Ames is currently working on the integration of the system to fit the thrusters and the PPU inside a 1.5U CubeSat together with the PhoneSat bus. This satellite is intended to be deployed from the ISS in 2015 and test the functionality of the thrusters by spinning the satellite around its long axis and measure the rotational speed with the phone gyros. This test flight will raise the TRL of the propulsion system from 5 to 7 and will be a first test for further CubeSats with propulsion systems, a key subsystem for long duration or interplanetary small satellite missions.

  19. On the Application of Hall Thruster Working with Ambient Atmospheric Gas for Orbital Station-Keeping

    Directory of Open Access Journals (Sweden)

    D. V. Duhopel'nikov

    2016-01-01

    Full Text Available The paper considers the application of the Hall thruster using the ambient atmospheric air for orbital station keeping. This is a relevant direction at the up-to-date development stage of propulsion systems. Many teams of designers of electric rocket thrusters evaluate the application of different schemes of particle acceleration at the low-earth orbit. Such technical solution allows us to abandon the storage systems of the working agent on the spacecraft board. Thus, lifetime of such a system at the orbit wouldn`t be limited by fuel range. The paper suggests a scheme of the propulsion device with a parabolic confuser that provides a required compression ratio of the ambient air for correct operation. Formulates physical and structural restrictions on ambient air to be used as a working agent of the thruster. Pointes out that the altitudes from 200 to 300 km are the most promising for such propulsion devices. Shows that for operation at lower altitudes are required the higher capacities that are not provided by modern onboard power supply systems. For the orbit heightening the air intakes with significant compression rate are of necessity. The size of such air intakes would exceed nose fairing of exploited space launch systems. To perform further design calculations are shown dependencies that allow us to calculate an effective diameter of the thruster channel and a critical voltage to be desirable for thrust force excess over air resistance. The dependencies to calculate minimal and maximal fluxes of neutral particles of oxygen and nitrogen, that are necessary for normal thruster operation, are also shown. Calculation results of the propulsion system parameters for the spacecrafts with cross-sectional area within 1 - 3 m2 and inlet diameter of air intake within 1 - 3 m are demonstrated. The research results have practical significance in design of advanced propulsion devices for lowaltitude spacecrafts. The work has been supported by the RFFR

  20. The electrodeless Lorentz force thruster experiment

    Science.gov (United States)

    Weber, Thomas E.

    The Electrodeless Lorentz Force (ELF) thruster is a novel type of plasma thruster, which utilizes Rotating Magnetic Field current drive within a diverging magnetic field to form, accelerate, and eject a Field Reversed Configuration plasmoid. The ELF program is a result of a Small Business Technology Transfer grant awarded to MSNW LLC by the Air Force Office of Scientific Research for the research of the revolutionary space propulsion concept represented by ELF. These grants are awarded to small businesses working in collaboration with a university, in this case, the University of Washington. The program was split into two concurrent research efforts; a numerical modeling study undertaken at the UW branch of the Plasma Science and Innovation Center, and an experimental effort taking place at the UW Plasma Dynamics Laboratory with additional support from MSNW (the latter being the subject of this dissertation). It is the aim of this dissertation is to present to the reader the necessary background information needed to understand the operation of the ELF thruster, an overview of the experimental setup, a review of the significant experimental findings, and a discussion regarding the operation and performance of the thruster.

  1. Hall Effect Thruster Ground Testing Challenges

    Science.gov (United States)

    2009-08-18

    conditional stability of the inverted pendulum thrust stand provides improved measurement sensitivity.5 With the displacement of the inverted pendulum...July 2005. 12Samiento, C., “RHETT2/ EPDM Hall Thruster Propulsion System Electromagnetic Compatability Evaluation,” Proceed- ings of the 25th

  2. Microdischarge plasma thrusters for small satellite propulsion

    Science.gov (United States)

    Raja, Laxminarayan

    2009-10-01

    Small satellites weighing less than 100 kg are gaining importance in the defense and commercial satellite community owing to advantages of low costs to build and operate, simplicity of design, rapid integration and testing, formation flying, and multi-vehicle operations. The principal challenge in the design and development of small satellite subsystems is the severe mass, volume, and power constraints posed by the overall size of the satellite. The propulsion system in particular is hard to down scale and as such poses a major stumbling block for small satellite technology. Microdischarge-based miniaturized plasma thrusters are potentially a novel solution to this problem. In its most basic form a microdischarge plasma thruster is a simple extension of a cold gas micronozzle propulsion device, where a direct or alternating current microdischarge is used to preheat the gas stream to improve to specific impulse of the device. We study a prototypical thruster device using a detailed, self-consistent coupled plasma and fluid flow computational model. The model describes the microdischarge power deposition, plasma dynamics, gas-phase chemical kinetics, coupling of the plasma phenomena with high-speed flow, and overall propulsion system performance. Unique computational challenges associated with microdischarge modeling in the presence of high-speed flows are addressed. Compared to a cold gas micronozzle, a significant increase in specific impulse (50 to 100 %) is obtained from the power deposition in the diverging supersonic section of the thruster nozzle. The microdischarge remains mostly confined inside the micronozzle and operates in an abnormal glow discharge regime. Gas heating, primarily due to ion Joule heating, is found to have a strong influence on the overall discharge behavior. The study provides a validation of the concept as simple and effective approach to realizing a relatively high-specific impulse thruster device at small geometric scales.

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

    Science.gov (United States)

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

    2012-01-01

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

  4. 纳秒紫外重复脉冲激光烧蚀单晶硅的热力学过程研究%Investigation of thermodynamic progress of silicon ablated by nanosecond uv repetitive pulse laser

    Institute of Scientific and Technical Information of China (English)

    包凌东; 韩敬华; 段涛; 孙年春; 高翔; 冯国英; 杨李茗; 牛瑞华; 刘全喜

    2012-01-01

    采用波长为355nm的纳秒紫外重复脉冲激光对单晶硅片进行了盲孔加工实验,观测了随脉冲增加激光烧蚀硅片的外观形貌和盲孔孔深、孔径的变化规律,并对紫外激光辐照硅片的热力学过程进行了分析.研究结果表明:紫外激光加工硅盲孔是基于热、力效应共同作用的结果,热效应会使得硅材料熔化、气化甚至发生电离产生激光等离子体,为材料的去除提供条件;激光等离子体冲击波以及高温气态物向外膨胀会对熔化材料产生压力致使其向外喷射,为重复脉冲的进一步烧蚀提供了条件;力效应主要沿着激光传输的方向,垂直于硅表面,使得去除部位主要集中在孔的深度方向,达到较高的孔径比,实验观察孔径比可达8:1;此外,激光等离子体的产生也阻止了激光对靶面的作用,加之随孔深的增加激光发生散焦,使得烧蚀深度有一定的限制,实验观察烧蚀脉冲个数在前100个时加工效率较高.%The blind holes processing experiment is conducted on the silicon under the radiation of a 355 nm nanosecond UV repetitive pulse laser. With the increase of the laser pulse number, the variations of the silicon morphology, the depth and aperture of the blind holes are observed, and the thermodynamic process of UV laser irradiating silicon is analyzed. The results show that the formation of the blind silicon hole in the laser ablation process is due to the interaction between thermal effect and force effect. Thermal effect results in fusion, vaporization and even producing laser plasma by ionization in silicon, which is essential to the removal of the material. The molten material is compressed by the plasma shock wave and the expansion of the high-temperature gaseous material,and then ejection outward, which will benefit the further ablation; the force propagates along the laser transmission direction,perpendicular to the silicon

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

    Science.gov (United States)

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

    2006-01-01

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

  6. A Coupled MHD and Thermal Model Including Electrostatic Sheath for Magnetoplasmadynamic Thruster Simulation

    Science.gov (United States)

    Kawasaki, Akira; Kubota, Kenichi; Funaki, Ikkoh; Okuno, Yoshihiro

    2016-09-01

    Steady-state and self-field magnetoplasmadynamic (MPD) thruster, which utilizes high-intensity direct-current (DC) discharge, is one of the prospective candidates of future high-power electric propulsion devices. In order to accurately assess the thrust performance and the electrode temperature, input electric power and wall heat flux must correctly be evaluated where electrostatic sheaths formed in close proximity of the electrodes affect these quantities. Conventional model simulates only plasma flows occurring in MPD thrusters with the absence of electrostatic sheath consideration. Therefore, this study extends the conventional model to a coupled magnetohydrodynamic (MHD) and thermal model by incorporating the phenomena relevant to the electrostatic sheaths. The sheaths are implemented as boundary condition of the MHD model on the walls. This model simulated the operation of the 100-kW-class thruster at discharge current ranging from 6 to 10 kA with argon propellant. The extended model reproduced the discharge voltages and wall heat load which are consistent with past experimental results. In addition, the simulation results indicated that cathode sheath voltages account for approximately 5-7 V subject to approximately 20 V of discharge voltages applied between the electrodes. This work was supported by JSPS KAKENHI Grant Numbers 26289328 and 15J10821.

  7. Development Status of Power Processing Unit for 250mN-Class Hall Thruster

    Science.gov (United States)

    Osuga, H.; Suzuki, K.; Ozaki, T.; Nakagawa, T.; Suga, I.; Tamida, T.; Akuzawa, Y.; Suzuki, H.; Soga, Y.; Furuichi, T.; Maki, S.; Matui, K.

    2008-09-01

    Institute for Unmanned Space Experiment Free Flyer (USEF) and Mitsubishi Electric Corporation (MELCO) are developing the next generation ion engine system under the sponsorship of Ministry of Economy, Trade and Industry (METI) within six years. The system requirement specifications are a thrust level of over 250mN and specific impulse of over 1500 sec with a less than 5kW electric power supply, and a lifetime of over 3,000 hours. These target specifications required the development of both a Hall Thruster and a Power Processing Unit (PPU). In the 2007 fiscal year, the PPU called Second Engineering Model (EM2) consist of all power supplies was a model for the Hall Thruster system. The EM2 PPU showed the discharge efficiency was over 96.2% for 250V and 350V at output power between 1.8kW to 4.5kW. And also the Hall Thruster could start up quickly and smoothly to control the discharge voltage, the inner magnet current, the outer magnet current and the xenon flow rate. This paper reports on the design and test results of the EM2 PPU.

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

    Science.gov (United States)

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

    2014-01-01

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

  9. Laser ablation in a running hall effect thruster for space propulsion

    Science.gov (United States)

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

    2013-07-01

    Hall Effect Thrusters (HETs) are promising electric propulsion devices for the station-keeping of geostationary satellites (more than 120 in orbit to date). Moreover, they can offer a cost-effective solution for interplanetary journey, as proved by the recent ESA SMART-1 mission to the Moon. The main limiting factor of the HETs lifetime is the erosion of the annular channel ceramics walls. In order to provide a better understanding of the energy deposition on the insulated walls, a laser irradiation study has been carried out on a PPS100-ML thruster during its run in the PIVOINE-2G ground test facility (CNRS Orléans, France). Two distinct approaches have been followed: continuous wave fiber laser irradiation (generation of thermal defects) and nanosecond pulsed laser ablation (generation of topological defects). The irradiated zones have been monitored in situ by IR thermography and optical emission spectroscopy and further investigated ex situ by scanning electron microscopy and profilometry.

  10. High-Pressure Lightweight Thrusters

    Science.gov (United States)

    Holmes, Richard; McKechnie, Timothy; Shchetkovskiy, Anatoliy; Smirnov, Alexander

    2013-01-01

    interface realizes pseudo-plastic behavior with significant increase in the tensile strength. The investigation of high-temperature strength of C/Cs under high-rate heating (critical for thrust chambers) shows that tensile and compression strength increases from 70 MPa at room temperature to 110 MPa at 1,773 K, and up to 125 MPa at 2,473 K. Despite these unique properties, the use of C/Cs is limited by its high oxidation rate at elevated temperatures. Lining carbon/carbon chambers with a thin layer of iridium or iridium and rhenium is an innovative way to use proven refractory metals and provide the oxidation barrier necessary to enable the use of carbon/ carbon composites. Due to the lower density of C/Cs as compared to SiC/SiC composites, an iridium liner can be added to the C/C structure and still be below the overall thruster weight. Weight calculations show that C/C, C/C with 50 microns of Ir, and C/C with 100 microns of Ir are of less weight than alternative materials for the same construction.

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

    Science.gov (United States)

    Schnelker, D. E.; Kami, S.

    1972-01-01

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

  12. Wide Output Range Power Processing Unit for Electric Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Hall thrusters can be operated over a wide range of specific impulse while maintaining high efficiency. However S/C power system constraints on electric propulsion...

  13. High Throughput 600 Watt Hall Effect Thruster for Space Exploration

    Science.gov (United States)

    Szabo, James; Pote, Bruce; Tedrake, Rachel; Paintal, Surjeet; Byrne, Lawrence; Hruby, Vlad; Kamhawi, Hani; Smith, Tim

    2016-01-01

    A nominal 600-Watt Hall Effect Thruster was developed to propel unmanned space vehicles. Both xenon and iodine compatible versions were demonstrated. With xenon, peak measured thruster efficiency is 46-48% at 600-W, with specific impulse from 1400 s to 1700 s. Evolution of the thruster channel due to ion erosion was predicted through numerical models and calibrated with experimental measurements. Estimated xenon throughput is greater than 100 kg. The thruster is well sized for satellite station keeping and orbit maneuvering, either by itself or within a cluster.

  14. High Efficiency Hall Thruster Discharge Power Converter Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek leveraged previous, internally sponsored, high power, Hall thruster discharge converter development which allowed it to design, build, and test new printed...

  15. Studies of Non-Conventional Configuration Closed Electron Drift Thrusters

    Energy Technology Data Exchange (ETDEWEB)

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

    2001-09-10

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

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

    Science.gov (United States)

    Shastry, Rohit; Huang, Wensheng; Kamhawi, Hani

    2015-01-01

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

  17. Thermal Characterization of a Hall Effect Thruster

    Science.gov (United States)

    2008-03-01

    Material Curie Temperature Iron 770 °C Nickel 358 °C Cobalt 1130 °C Gadolinium 20 °C Terfenol 380-430 °C Alnico 850 °C Hard Ferrites 400-700...C Barium Ferrite 450 °C Hall Effect thrusters generally use iron magnets with a Curie temperature of 770 °C. Decreasing the magnetic strength

  18. ELECTROSTATIC ION THRUSTERS - TOWARDS PREDICTIVE MODELING

    Directory of Open Access Journals (Sweden)

    Julia Duras

    2015-02-01

    Full Text Available For satellite missions, thrusters have to be qualified in large vacuum vessels to simulate space environment. One caveat of these experiments is the possible  modification of the beam properties due to the interaction of the energetic ions with the  vessel walls. Impinging ions can produce sputtered impurities or secondary  electrons from the wall. These can stream back into the acceleration channel of the  thruster and produce co-deposited layers. Over the long operation time of thousands  of hours, such layers can modify the optimized geometry and induce changes of the ion beam properties, e.g. broadening of the angular distribution and thrust reduction. To study such effects, a Monte Carlo code for the simulation of the interaction of ion thruster beams with vessel  walls was developed. Strategies to overcome sputter limitations by additional baffles are  studied with the help of this Monte-Carlo erosion code.

  19. Multi-Scale Modeling of Plasma Thrusters

    Science.gov (United States)

    Batishchev, Oleg

    2004-11-01

    Plasma thrusters are characterized with multiple spatial and temporal scales, which are due to the intrinsic physical processes such as gas ionization, wall effects and plasma acceleration. Characteristic times for hot plasma and cold gas are differing by 6-7 orders of magnitude. The typical collisional mean-free-paths vary by 3-5 orders along the devices. These make questionable a true self-consistent modeling of the thrusters. The latter is vital to the understanding of complex physics, non-linear dynamics and optimization of the performance. To overcome this problem we propose the following approach. All processes are divided into two groups: fast and slow. The slow ones include gas evolution with known sources and ionization sink. The ionization rate, transport coefficients, energy sources are defined during "fast step". Both processes are linked through external iterations. Multiple spatial scales are handled using moving adaptive mesh. Development and application of this method to the VASIMR helicon plasma source and other thrusters will be discussed. Supported by NASA.

  20. Plume Comparisons between Segmented Channel Hall Thrusters

    Science.gov (United States)

    Niemack, Michael; Staack, David; Raitses, Yevgeny; Fisch, Nathaniel

    2001-10-01

    Angular ion flux plume measurements were taken in several configurations of segmented channel Hall thrusters. The configurations differed by the placement of relatively short rings made from materials with different conductive and secondary electron emission properties along the boron nitride ceramic channel of the thrusters (these have been shown to affect the plume [1]). The ion fluxes are compared with ion trajectory simulations based on plasma potential data acquired with a high speed emissive probe [2]. Preliminary results indicate that in addition to the physical properties of the segments, the plume angle can be strongly affected by the placement of segmented rings relative to the external and internal walls of the channel. [1] Y. Raitses, L. Dorf, A. Litvak and N. J. Fisch, Journal of Applied Physics 88, 1263, 2000 [2] D. Staack, Y. Raitses, N. J. Fisch, Parametric Investigations of Langmuir Probe Induced Perturbations in a Hall Thruster, DPP01 Poster Presentation This work was supported by the U.S. DOE Contract No. DE-ACO2-76-CHO3073.

  1. Externally Wetted Ionic Liquid Thruster

    Science.gov (United States)

    Lozano, P.; Martinez-Sanchez, M.; Lopez-Urdiales, J. M.

    2004-10-01

    This paper presents initial developments of an electric propulsion system based on ionic liquid ion sources (ILIS). Propellants are ionic liquids, which are organic salts with two important characteristics; they remain in the liquid state at room temperature and have negligible vapor pressure, thus allowing their use in vacuum. The working principles of ILIS are similar to those of liquid metal ion sources (LMIS), in which a Taylor cone is electrostatically formed at the tip of an externally wetted needle while ions are emitted directly from its apex. ILIS have the advantage of being able to produce negative ions that have similar masses than their positive counterparts with similar current levels. This opens up the possibility of achieving plume electrical neutrality without electron emitters. The possible multiplexing of these emitters is discussed in terms of achievable thrust density for applications other than micro-propulsion.

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

    Science.gov (United States)

    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

  3. Effect of Segmented Electrode Length on the Performances of an Aton-Type Hall Thruster

    Institute of Scientific and Technical Information of China (English)

    DUAN Ping; BIAN Xingyu; CAO Anning; LIU Guangrui; CHEN Long; YIN Yan

    2016-01-01

    The influences of the low-emissive graphite segmented electrode placed near the channel exit on the discharge characteristics of a Hall thruster are studied using the particlein-cell method.A two-dimensional physical model is established according to the Hall thruster discharge channel configuration.The effects of electrode length on the potential,ion density,electron temperature,ionization rate and discharge current are investigated.It is found that,with the increasing of the segmented electrode length,the equipotential lines bend towards the channel exit,and approximately parallel to the wall at the channel surface,the radial velocity and radial flow of ions are increased,and the electron temperature is also enhanced.Due to the conductive characteristic of electrodes,the radial electric field and the axial electron conductivity near the wall are enhanced,and the probability of the electron-atom ionization is reduced,which leads to the degradation of the ionization rate in the discharge channel.However,the interaction between electrons and the wall enhances the near wall conductivity,therefore the discharge current grows along with the segmented electrode length,and the performance of the thruster is also affected.

  4. Electrostatic Probe with Shielded Probe Insulator Tube for Low Disturbing Plasma Measurements in Hall Thrusters

    Energy Technology Data Exchange (ETDEWEB)

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

    2003-07-10

    Electrostatic probes are widely used to measure spatial plasma parameters of the quasi-neutral plasma in Hall thrusters and similar ExB electric discharge devices. Significant perturbations of the plasma, induced by such probes, can mask the actual physics involved in operation of these devices. In Hall thrusters, probe-induced perturbations can produce changes in the discharge current and plasma parameters on the order of their steady state values. These perturbations are explored by varying the material, penetration distance, and residence time of various probe designs. A possible cause of these perturbations appears to be the secondary electron emission, induced by energetic plasma electrons, from insulator ceramic tubes in which the probe wire is inserted. A new probe in which a low secondary electron emission material, such as metal, shields the probe ceramic tube, is shown to function without producing such large perturbations. A segmentation of this shield further prevents probe -induced perturbations, by not shortening the plasma through the conductive shield. In a set of experiments with a segmented shield probe, the thruster was operated in the input power range of 500-2.5 kW and discharge voltages of 200-500 V, while the probe-induced perturbations of the discharge current were below 4% of its steady state value in the region in which 90% of the voltage drop takes place.

  5. An Investigation into the Spectral Imaging of Hall Thruster Plumes

    Science.gov (United States)

    2015-07-01

    zone shifting to a more upstream location in the discharge channel as observed in Ref. 14 for the BHT -600 thruster likely due to increased electron...to a more upstream location in the discharge channel as observed in Ref. 14 for the BHT -600 thruster likely due to increased electron mobility from

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2008-04-23

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

  7. Feasibility assessment of magnetic sensors for measurment of Hall current induced changes to the static magnetic field nearby a Hall thruster

    Science.gov (United States)

    Morozko, Zoe

    A Hall thruster is an electric propulsion device that produces thrust electrostatically by accelerating propellant to velocities 5 to 10 times higher than is achievable using conventional chemical thrusters. This is accomplished through the application of static, crossed electric and magnetic fields that are concentrated in a region close to the exit plane of the thruster. During operation an azimuthal plasma-electron current develops in the region where the electric and magnetic fields are concentrated. This embedded plasma current is referred to as the Hall current. The thrust produced from accelerating the propellant is transferred to a satellite or spacecraft through interaction between the Hall current and the magnetic coils used to produce the static magnetic field within the thruster. The Hall current can be calculated and the thrust can be determined in real time by measuring the magnetic field produced by the Hall current using sensors located external to the thruster. This work investigates the feasibility of placing magnetic sensors in the regions close to the exit of the thruster to measure the external magnetic field and correlate it to the Hall current. A finite element magnetic solver was used to identify several locations outside of the thrust plume and near the pole piece where the magnetic field magnitude changes by several Gauss in a background field level of ˜50 Gauss. Magnetic sensors based on the giant magnetoresistive effect were identified as acceptable with regard to sensitivity, and measurements made with these sensors in a simulated high background magnetic field environment demonstrated that changes of 0.5 Gauss could be easily measured. This work also presents the development of a thrust stand that will be useful in future work to demonstrate the overall concept. Special focus was directed to the design of the data acquisition system and in-vacuum calibration system used to make measurements with the thrust stand.

  8. Recent activities in the development of the MOA thruster

    Science.gov (United States)

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

    2008-07-01

    More than 60 years after the later Nobel laureate Hannes Alfvén 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 Alfvén waves for propulsive purposes has been proposed, patented and examined for the first time by a group of inventors. The name of the concept, utilising Alfvén waves to accelerate ionised matter for propulsive purposes, is MOA-magnetic field oscillating amplified thruster. Alfvén waves are generated by making use of two coils, one being permanently powered and serving also as magnetic nozzle, the other one being switched on and off in a cyclic way, deforming the field lines of the overall system. It is this deformation that generates Alfvén waves, which are in the next step used to transport and compress the propulsive medium, in theory leading to a propulsion system with a much higher performance than any other electric propulsion system. Based on computer simulations, which were conducted to get a first estimate on the performance of the system, MOA is a corrosion free and highly flexible propulsion system, whose performance parameters might easily be adapted in flight, by changing the mass flow and/or the power level. As such the system is capable to deliver a maximum specific impulse of 13 116 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. First tests-that are further described in this paper-have been conducted successfully and underline the feasibility of the concept. 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

  9. Azimuthal Spoke Propagation in Hall Effect Thrusters

    Science.gov (United States)

    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.

  10. Micro-Discharge Micro-Thruster

    Science.gov (United States)

    2005-06-01

    breakdown at the maximum applied voltage (900 V) in Argon. The back side of the Paschen curve for Ar occurs at a pressure-length (P·d) product of less than...significant capacitance to ground from either lead (~ 100 nF). As small as this is, it had a profound effect on the discharge (see next section). A more space... effect in most thrusters even in the 100 Watt class. For a micro-discharge, even a stray coupling capacitance 50 pF observed for the power leads

  11. Electric propulsion for satellites and spacecraft: established technologies and novel approaches

    Science.gov (United States)

    Mazouffre, Stéphane

    2016-06-01

    This contribution presents a short review of electric propulsion (EP) technologies for satellites and spacecraft. Electric thrusters, also termed ion or plasma thrusters, deliver a low thrust level compared to their chemical counterparts, but they offer significant advantages for in-space propulsion as energy is uncoupled to the propellant, therefore allowing for large energy densities. Although the development of EP goes back to the 1960s, the technology potential has just begun to be fully exploited because of the increase in the available power aboard spacecraft, as demonstrated by the very recent appearance of all-electric communication satellites. This article first describes the fundamentals of EP: momentum conservation and the ideal rocket equation, specific impulse and thrust, figures of merit and a comparison with chemical propulsion. Subsequently, the influence of the power source type and characteristics on the mission profile is discussed. Plasma thrusters are classically grouped into three categories according to the thrust generation process: electrothermal, electrostatic and electromagnetic devices. The three groups, along with the associated plasma discharge and energy transfer mechanisms, are presented via a discussion of long-standing technologies like arcjet thrusters, magnetoplasmadynamic thrusters, pulsed plasma thrusters and ion engines, as well as Hall thrusters and variants. More advanced concepts and new approaches for performance improvement are discussed afterwards: magnetic shielding and wall-less configurations, negative ion thrusters and plasma acceleration with a magnetic nozzle. Finally, various alternative propellant options are analyzed and possible research paths for the near future are examined.

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

    Science.gov (United States)

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

    2017-03-01

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

  13. Development of High-Power Hall Thruster Power Processing Units at NASA GRC

    Science.gov (United States)

    Pinero, Luis R.; Bozak, Karin E.; Santiago, Walter; Scheidegger, Robert J.; Birchenough, Arthur G.

    2015-01-01

    NASA GRC successfully designed, built and tested four different power processor concepts for high power Hall thrusters. Each design satisfies unique goals including the evaluation of a novel silicon carbide semiconductor technology, validation of innovative circuits to overcome the problems with high input voltage converter design, development of a direct-drive unit to demonstrate potential benefits, or simply identification of lessonslearned from the development of a PPU using a conventional design approach. Any of these designs could be developed further to satisfy NASA's needs for high power electric propulsion in the near future.

  14. Two-Dimensional, Time-Dependent Plasma Structures of a Hall Effect Thruster

    Science.gov (United States)

    2011-09-01

    47 3.7 Relative spectral response of the Shimadzu HPV -2 ultra-high speed camera taken from Shimadzu HPV -2 Spectral Response . 48 3.8 Sample...a National Instruments SCXI-1321 to measure the current of each component in real-time. 41 Figure 3.4: Electrical diagram of Hall thruster components...experiment. The direct emission data was collected with a Shimadzu HyperVision HPV -2 high-speed camera which is able to record at one-million frames per

  15. Cassini Thruster Calibration Algorithm Using Reaction Wheel Biasing Data

    Science.gov (United States)

    Rizvi, Farheen

    2012-01-01

    Thrust force estimates for the reaction control thrusters on-board Cassini spacecraft are presented in this paper. Cassini consists of two thruster branches (A and B) each with eight thrusters. The four Z-thrusters control the X and Y-axes, while the four Y-thrusters control the Z-axis. It is important to track the thrust force estimates in order to detect any thruster degradation and for supporting various activities in spacecraft operations (Titan flyby, spacecraft maneuvers). The Euler equation, which describes the rotational motion of the spacecraft during a reaction wheel bias event, is used to develop the algorithm. The thrust estimates are obtained from the pseudo inverse solution using flight telemetry during the bias. Results show that the A-branch Z3A and Z4A thrusters exhibited degraded thrust in November 2008. Due to the degraded thrust performance of Z3A and Z4A, A-branch usage was discontinued and prime branch was swapped to B-branch in March 2009. The thrust estimates from the B-branch do not show any degradation to date. The algorithm is used to trend the B-branch thrust force estimates as the mission continues.

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

    Science.gov (United States)

    Peterson, Peter Y.

    2015-01-01

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

  17. Transient tests on an MHD thruster

    Energy Technology Data Exchange (ETDEWEB)

    Pierson, E.S. (Purdue Univ., Hammond, IN (United States). Dept. of Engineering); Libera, J.; Petrick, M. (Argonne National Lab., IL (United States). Energy Systems Div.)

    1993-01-01

    Three different types of transient tests were made -- coast downs to zero voltage and current under open circuit and short circuit conditions, reverses where the applied voltage was reversed to the same or a different value, and jumps where the voltage applied to the thruster was increased without a change in polarity. Most except the coast downs were dons both quickly (voltage changes as fast as possible) and slowly (6 s to complete the voltage change). A few slower (12 s) transients were done. Transient runs were made for water conductivities of 16.2 and 5.09 S/m. In all cases steady-state conditions were established and several seconds of data taken before initiating the transients. Data were measured every 0.75 to 1 .5 second over the time interval of interest. Particular attention was paid to looking for evidence of gas bubbles, and to the chance of the voltage profiles between the electrodes. The data are interpreted based on the behavior of the power supply and the thruster.

  18. Optimisation of a quantum pair space thruster

    Directory of Open Access Journals (Sweden)

    Valeriu DRAGAN

    2012-06-01

    Full Text Available The paper addresses the problem of propulsion for long term space missions. Traditionally a space propulsion unit has a propellant mass which is ejected trough a nozzle to generate thrust; this is also the case with inert gases energized by an on-board power unit. Unconventional methods for propulsion include high energy LASERs that rely on the momentum of photons to generate thrust. Anti-matter has also been proposed for energy storage. Although the momentum of ejected gas is significantly higher, the LASER propulsion offers the perspective of unlimited operational time – provided there is a power source. The paper will propose the use of the quantum pair formation for generating a working mass, this is different than conventional anti-matter thrusters since the material particles generated are used as propellant not as energy storage.Two methods will be compared: LASER and positron-electron, quantum pair formation. The latter will be shown to offer better momentum above certain energy levels.For the demonstrations an analytical solution is obtained and provided in the form of various coefficients. The implications are, for now, theoretical however the practicality of an optimized thruster using such particles is not to be neglected for long term space missions.

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

    Science.gov (United States)

    Low, C. A., Jr.

    1975-01-01

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

  20. A low power pulsed arcjet thruster for spacecraft propulsion

    Science.gov (United States)

    Willmes, Gary Francis

    1997-11-01

    An electrothermal thruster that operates in a pulsed mode at low power (pendulum-type thrust stand, and input power levels from 24 to 119 watts are determined from measurements of pulse rate and breakdown voltage. A maximum specific impulse of 305 seconds is achieved with 38% efficiency. A time-dependent, quasi-1D numerical model is developed to evaluate energy losses in the pulsed arcjet. The numerical model uses a time-marching procedure and the MacCormack predictor-corrector algorithm. Viscous and heat transfer effects are incorporated though a friction factor and an average heat transfer coefficient. A numerical study of nozzle parameters, capillary geometry, wall temperature, and pulse energy shows that the performance is insensitive to capillary and nozzle geometry and that thermal characteristics are the dominant factor affecting performance. The specific impulse and efficiency of the pulsed arcjet are found to be sensitive to wall temperature due to heat transfer losses in the subsonic region. A pulse-forming electrical circuit is developed to reduce energy losses in the storage capacitor, and greater than 85% of the initial stored energy is transferred to the arc in a unipolar pulse. A high current diode installed across the capacitor terminals is used to eliminate voltage reversals in the current. The experimental breakdown voltage of the helium gas between the electrodes is found to follow a Paschen relationship where the minimum electrode separation distance is used in evaluating the data.

  1. Characteristics of plasma properties in an ablative pulsed plasma thruster

    Energy Technology Data Exchange (ETDEWEB)

    Schoenherr, Tony; Nees, Frank; Arakawa, Yoshihiro [Department of Aeronautics and Astronautics, University of Tokyo, Bunkyo, Tokyo 113-8656 (Japan); Komurasaki, Kimiya [Department of Advanced Energy, University of Tokyo, Kashiwa, Chiba 277-8561 (Japan); Herdrich, Georg [Institute of Space Systems (IRS), University of Stuttgart, 70569 Stuttgart, Baden-Wuerttemberg (Germany)

    2013-03-15

    Pulsed plasma thrusters are electric space propulsion devices which create a highly transient plasma bulk in a short-time arc discharge that is expelled to create thrust. The transitional character and the dependency on the discharge properties are yet to be elucidated. In this study, optical emission spectroscopy and Mach-Zehnder interferometry are applied to investigate the plasma properties in variation of time, space, and discharge energy. Electron temperature, electron density, and Knudsen numbers are derived for the plasma bulk and discussed. Temperatures were found to be in the order of 1.7 to 3.1 eV, whereas electron densities showed maximum values of more than 10{sup 17} cm{sup -3}. Both values showed strong dependency on the discharge voltage and were typically higher closer to the electrodes. Capacitance and time showed less influence. Knudsen numbers were derived to be in the order of 10{sup -3}-10{sup -2}, thus, indicating a continuum flow behavior in the main plasma bulk.

  2. Kinetic models for the VASIMR thruster helicon plasma source

    Science.gov (United States)

    Batishchev, Oleg; Molvig, Kim

    2001-10-01

    Helicon gas discharge [1] is widely used by industry because of its remarkable efficiency [2]. High energy and fuel efficiencies make it very attractive for space electrical propulsion applications. For example, helicon plasma source is used in the high specific impulse VASIMR [3] plasma thruster, including experimental prototypes VX-3 and upgraded VX-10 [4] configurations, which operate with hydrogen (deuterium) and helium plasmas. We have developed a set of models for the VASIMR helicon discharge. Firstly, we use zero-dimensional energy and mass balance equations to characterize partially ionized gas condition/composition. Next, we couple it to one-dimensional hybrid model [6] for gas flow in the quartz tube of the helicon. We compare hybrid model results to a purely kinetic simulation of propellant flow in gas feed + helicon source subsystem. Some of the experimental data [3-4] are explained. Lastly, we discuss full-scale kinetic modeling of coupled gas and plasmas [5-6] in the helicon discharge. [1] M.A.Lieberman, A.J.Lihtenberg, 'Principles of ..', Wiley, 1994; [2] F.F.Chen, Plas. Phys. Contr. Fus. 33, 339, 1991; [3] F.Chang-Diaz et al, Bull. APS 45 (7) 129, 2000; [4] J.Squire et al., Bull. APS 45 (7) 130, 2000; [5] O.Batishchev et al, J. Plasma Phys. 61, part II, 347, 1999; [6] O.Batishchev, K.Molvig, AIAA technical paper 2000-3754, -14p, 2001.

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

    Energy Technology Data Exchange (ETDEWEB)

    West, Michael D.; Charles, Christine; Boswell, Rod W. [Space Plasma, Power and Propulsion Group, Research School of Physics and Engineering, Australian National University, Canberra ACT 0200 (Australia)

    2009-05-15

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

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

    Science.gov (United States)

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

    2009-05-01

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

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

    Science.gov (United States)

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

    2017-02-01

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

  6. Self consistent kinetic simulations of SPT and HEMP thrusters including the near-field plume region

    CERN Document Server

    Matyash, K; Mutzke, A; Kalentev, O; Taccogna, F; Koch, N; Schirra, M

    2009-01-01

    The Particle-in-Cell (PIC) method was used to study two different ion thruster concepts - Stationary Plasma Thrusters (SPT) and High Efficiency Multistage Plasma Thrusters (HEMP-T), in particular the plasma properties in the discharge chamber due to the different magnetic field configurations. Special attention was paid to the simulation of plasma particle fluxes on the thrusters channel surfaces. In both cases, PIC proved itself as a powerful tool, delivering important insight into the basic physics of the different thruster concepts. The simulations demonstrated that the new HEMP thruster concept allows for a high thermal efficiency due to both minimal energy dissipation and high acceleration efficiency. In the HEMP thruster the plasma contact to the wall is limited only to very small areas of the magnetic field cusps, which results in much smaller ion energy flux to the thruster channel surface as compared to SPT. The erosion yields for dielectric discharge channel walls of SPT and HEMP thrusters were calc...

  7. 20mN, Variable Specific Impulse Colloid Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — During Phase I, Busek designed and manufactured an electrospray emitter capable of generating 20 mN in a 7" x 7" x 1.7" package. The thruster consists of nine...

  8. 20mN, Variable Specific Impulse Colloid Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Colloid thrusters have long been known for their exceptional thrust efficiency and ability to operate over a range of specific impulse due to easily variable...

  9. A High Performance Cathode Heater for Hall Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — High current hollow cathodes are the baseline electron source for next generation high power Hall thrusters. Currently for electron sources providing current levels...

  10. High Throughput Hall Thruster for Small Spacecraft Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek Co. Inc. proposes to develop a high throughput, nominal 100 W Hall Effect Thruster (HET). This HET will be sized for small spacecraft (< 180 kg), including...

  11. Magnesium Hall Thruster for Solar System Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The innovation being developed in this program is a Mg Hall Effect Thruster system that would open the door for In-Situ Resource Utilization based solar system...

  12. Plume Characterization of Busek 600W Hall Thruster

    Science.gov (United States)

    2012-03-09

    Dr. William A. Hargus Jr. (Member) Date iv Abstract The BHT -600W thruster has a high potential to place on various commercial and...Thrust Measurement ........................................................................................71 A. BHT -200W...71 B. BHT -600W’s Performance

  13. Radio Frequency Micro Ion Thruster for Precision Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop radio frequency discharge, gridded micro-ion thruster that produces sub-mN thrust precisely adjustable over a wide dynamic thrust range....

  14. Radio Frequency Micro Ion Thruster for Precision Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to continue development of an engineering model radio frequency discharge, gridded micro ion thruster that produces sub-mN to mN thrust precisely...

  15. Pulsed Electrogasdynamic Thruster for Attitude Control and Orbit Maneuver Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In the Phase I program we successfully demonstrated the feasibility of the Pulsed ElectroGasdynamic (PEG) thruster for attitude control and orbital maneuvering. In...

  16. Thruster Modelling for Underwater Vehicle Using System Identification Method

    Directory of Open Access Journals (Sweden)

    Mohd Shahrieel Mohd Aras

    2013-05-01

    Full Text Available This paper describes a study of thruster modelling for a remotely operated underwater vehicle (ROV by system identification using Microbox 2000/2000C. Microbox 2000/2000C is an XPC target machine device to interface between an ROV thruster with the MATLAB 2009 software. In this project, a model of the thruster will be developed first so that the system identification toolbox in MATLAB can be used. This project also presents a comparison of mathematical and empirical modelling. The experiments were carried out by using a mini compressor as a dummy depth pressure applied to a pressure sensor. The thruster model will thrust and submerge until it reaches a set point and maintain the set point depth. The depth was based on pressure sensor measurement. A conventional proportional controller was used in this project and the results gathered justified its selection.

  17. High Input Voltage Hall Thruster Discharge Converter Project

    Data.gov (United States)

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

  18. High Throughput Hall Thruster for Small Spacecraft Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek is developing a high throughput nominal 100-W Hall Effect Thruster. This device is well sized for spacecraft ranging in size from several tens of kilograms to...

  19. Modeling Common Cause Failures of Thrusters on ISS Visiting Vehicles

    Science.gov (United States)

    Haught, Megan

    2014-01-01

    This paper discusses the methodology used to model common cause failures of thrusters on the International Space Station (ISS) Visiting Vehicles. The ISS Visiting Vehicles each have as many as 32 thrusters, whose redundancy makes them susceptible to common cause failures. The Global Alpha Model (as described in NUREG/CR-5485) can be used to represent the system common cause contribution, but NUREG/CR-5496 supplies global alpha parameters for groups only up to size six. Because of the large number of redundant thrusters on each vehicle, regression is used to determine parameter values for groups of size larger than six. An additional challenge is that Visiting Vehicle thruster failures must occur in specific combinations in order to fail the propulsion system; not all failure groups of a certain size are critical.

  20. Magnesium Hall Thruster for Solar System Exploration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to prove the feasibility of a Mg Hall effect thruster system that would open the door for In-Situ Resource Utilization (ISRU) based solar system...

  1. Long Life Cold Cathodes for Hall effect Thrusters Project

    Data.gov (United States)

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

  2. Thruster Modelling for Underwater Vehicle Using System Identification Method

    Directory of Open Access Journals (Sweden)

    Mohd Shahrieel Mohd Aras

    2013-05-01

    Full Text Available Abstract This paper describes a study of thruster modelling for a remotely operated underwater vehicle (ROV by system identification using Microbox 2000/2000C. Microbox 2000/2000C is an XPC target machine device to interface between an ROV thruster with the MATLAB 2009 software. In this project, a model of the thruster will be developed first so that the system identification toolbox in MATLAB can be used. This project also presents a comparison of mathematical and empirical modelling. The experiments were carried out by using a mini compressor as a dummy depth pressure applied to a pressure sensor. The thruster model will thrust and submerge until it reaches a set point and maintain the set point depth. The depth was based on pressure sensor measurement. A conventional proportional controller was used in this project and the results gathered justified its selection.

  3. Near-Term Laser Launch Capability: The Heat Exchanger Thruster

    Science.gov (United States)

    Kare, Jordin T.

    2003-05-01

    The heat exchanger (HX) thruster concept uses a lightweight (up to 1 MW/kg) flat-plate heat exchanger to couple laser energy into flowing hydrogen. Hot gas is exhausted via a conventional nozzle to generate thrust. The HX thruster has several advantages over ablative thrusters, including high efficiency, design flexibility, and operation with any type of laser. Operating the heat exchanger at a modest exhaust temperature, nominally 1000 C, allows it to be fabricated cheaply, while providing sufficient specific impulse (~600 seconds) for a single-stage vehicle to reach orbit with a useful payload; a nominal vehicle design is described. The HX thruster is also comparatively easy to develop and test, and offers an extremely promising route to near-term demonstration of laser launch.

  4. Four Thruster Microfluidic Electrospray Propulsion (MEP) Cubesat Board Demonstration Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Cubesat Microfluidic Electrospray Propulsion (MEP) system module prototype will be designed, built and tested to demonstrate that a four MEP thruster system can...

  5. Self-Healing Field-Emission Neutralizers for Electric Propulsion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Electric propulsion (EP) thrusters have the potential to enhance or enable Discovery-class missions. However, a significant challenge in scaling micro (< 100 W)...

  6. Monopropellant Thruster Development Using a Family of Micro Reactors

    Science.gov (United States)

    2017-02-17

    SCALE IN MILES HWY 395 ROSAMOND BLVD...AVENUE E H IG H W AY 1 4 LA N C A S TE R B LV D . 14 0t h S TR E E T E A S T RESERVATION BOUNDARY 0 5 10 SCALE IN MILES HWY 395 ROSAMOND BLVD...Monopropellant Thrusters Physical Description • Small (~1-22N) Thrusters Used for Attitude Control and Maneuvering of Small Spacecraft. AF-M315E

  7. Capillary Discharge Thruster Experiments and Modeling (Briefing Charts)

    Science.gov (United States)

    2016-06-01

    PROPULSION MODELS & EXPERIMENTS Spacecraft Propulsion Relevant Plasma: From hall thrusters to plumes and fluxes on components Complex reaction physics i.e...PROPULSION MODELS & EXPERIMENTS Spacecraft Propulsion Relevant Plasma: From hall thrusters to plumes and fluxes on components Complex reaction ...Conductivity h is the Enthalpy Cs is the Sound Speed Θ is the Wall Energy Flux Pekker, 40th AIAA Plasmadynamics and Laser Conference, 2009. R.S. MARTIN (ERC INC

  8. High-Efficiency Nested Hall Thrusters for Robotic Solar System Exploration

    Science.gov (United States)

    Hofer, Richard R.

    2013-01-01

    This work describes the scaling and design attributes of Nested Hall Thrusters (NHT) with extremely large operational envelopes, including a wide range of throttleability in power and specific impulse at high efficiency (>50%). NHTs have the potential to provide the game changing performance, powerprocessing capabilities, and cost effectiveness required to enable missions that cannot otherwise be accomplished. NHTs were first identified in the electric propulsion community as a path to 100- kW class thrusters for human missions. This study aimed to identify the performance capabilities NHTs can provide for NASA robotic and human missions, with an emphasis on 10-kW class thrusters well-suited for robotic exploration. A key outcome of this work has been the identification of NHTs as nearly constant-efficiency devices over large power throttling ratios, especially in direct-drive power systems. NHT systems sized for robotic solar system exploration are predicted to be capable of high-efficiency operation over nearly their entire power throttling range. A traditional Annular Hall Thruster (AHT) consists of a single annular discharge chamber where the propellant is ionized and accelerated. In an NHT, multiple annular channels are concentrically stacked. The channels can be operated in unison or individually depending on the available power or required performance. When throttling an AHT, performance must be sacrificed since a single channel cannot satisfy the diverse design attributes needed to maintain high thrust efficiency. NHTs can satisfy these requirements by varying which channels are operated and thereby offer significant benefits in terms of thruster performance, especially under deep power throttling conditions where the efficiency of an AHT suffers since a single channel can only operate efficiently (>50%) over a narrow power throttling ratio (3:1). Designs for 10-kW class NHTs were developed and compared with AHT systems. Power processing systems were

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2003-03-01

    A 2-kW Hall thruster was developed, built, and operated in an upgraded vacuum facility. The thruster performance and parameters of the plasma flow were measured by new diagnostics for plume measurements and plasma measurements inside the thruster channel. The thruster demonstrated efficient operation in terms of propellant and current utilization efficiencies in the input power range of 0.5-3.5 kW. Preliminary measurements of the ion energy spectra from the thruster axis region and the distribution of plasma parameters in the vicinity of the thruster exit are reported.

  10. Plasma plume diagnostics of low power stationary plasma thruster (SPT-20M8) with collisional radiative model

    Science.gov (United States)

    Uttamsing Rajput, Rajendrasing; Alona, Khaustova; Loyan, Andriy V.

    2017-03-01

    Electric propulsion offers higher specific impulse compared to the chemical propulsion systems. It reduces the overall propellant mass and enables high operational lifetimes. Scientific Technological Center of Space Power and Energy (STC SPE), KhAI is involved in designing, manufacturing and testing of stationary plasma thrusters (SPT). Efforts are made to perform plasma diagnostics with corona and collisional radiative models (C-R model), as expected corona model falls short below 4 eV because of the heavy particle collisions elimination, whereas the C-R model's applicability is confirmed. Several tests are performed to analyze the electron temperature at various operational parameters of thruster like discharge voltage and mass flow rate. SPT-20M8 far and near-field plumes diagnostics are performed. Feasibility of C-R model by comparing its result to optical emission spectroscopy (OES) to investigate the electron temperature is validated with the probe measurements within the 10% of discrepancy.

  11. Laser-Driven Mini-Thrusters

    Science.gov (United States)

    Sterling, Enrique; Lin, Jun; Sinko, John; Kodgis, Lisa; Porter, Simon; Pakhomov, Andrew V.; Larson, C. William; Mead, Franklin B.

    2006-05-01

    Laser-driven mini-thrusters were studied using Delrin® and PVC (Delrin® is a registered trademark of DuPont) as propellants. TEA CO2 laser (λ = 10.6 μm) was used as a driving laser. Coupling coefficients were deduced from two independent techniques: force-time curves measured with a piezoelectric sensor and ballistic pendulum. Time-resolved ICCD images of the expanding plasma and combustion products were analyzed in order to determine the main process that generates the thrust. The measurements were also performed in a nitrogen atmosphere in order to test the combustion effects on thrust. A pinhole transmission experiment was performed for the study of the cut-off time when the ablation/air breakdown plasma becomes opaque to the incoming laser pulse.

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

    Directory of Open Access Journals (Sweden)

    Anna-Maria Theodora ANDREESCU

    2016-12-01

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

  13. The X3: A 200 kW Class Nested Channel Hall Thruster

    Science.gov (United States)

    Sheehan, J. P.

    2016-10-01

    Electric propulsion has seen rapid adoption in recent years for commercial, scientific, and exploratory space missions. The X3 is a three channel nested channel Hall thruster, designed to push the boundaries of high power electric propulsion for cargo transfer to Mars and large military assets. It has been operated at thermal steady state up to 30 kW of power. Thrust measurements were made on an inverted pendulum thrust stand, indicating over 2000 s specific impulse and 65 mN/kW thrust to power ratio. Detailed plume measurements were made with Faraday and Langmuir probes. The multiple concentric channels provide better performance than the sum of the individual channel operations due to superior propellant utilization from its compact design. Using a high speed camera, the breathing and spoke mode instabilities were captured in all three channels. Spoke and breathing instabilities couple between the channels, indicating that complex plasma and neutral interactions are at play. Electron transport, both cross field and in the cathode plume, are well suited to be explored in a thruster of this size. Supported under NASA contract No. NNH16CP17C.

  14. Design and development of the Army KE ASAT ACS thruster

    Science.gov (United States)

    Craddock, Jeff; Janeski, Bruce

    1993-06-01

    Increasingly ambitious missions for advanced kinetic energy (KE) weapons have necessitated the development of a lightweight storable-propellant attitude control system (ACS) thruster capable of very fast response and long duration firings. This paper summarizes the results of a ACS thruster design and development test effort, performed for the U.S. Army Space and Strategic Defense Command (USASSDC) on the KE Anti Satellite (KE ASAT) weapon system program. Design approaches used to achieve long-duration continuous firing with a composite combustion chamber are detailed. This design effort culminated in a 6.7 lbf. thruster assembly weighing less than 0.2 pounds, approximately one-sixth that of a conventional satellite ACS thruster. Results of tests of flightweight engines with nitrogen tetroxide and monomethyl hydrazine hypergolic propellants are included. The test series culminated in what is believed to be the industry's longest continuous firing of a composite combustion chamber. This thruster will be integrated into the KE ASAT kinetic vehicle for its first free-flight hover test in early FY94. The demonstrated fast response, high pulse performance, and long-duration capabilities of this engine suggest that this thruster can significantly increase the capability of other spacecraft.

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

    Energy Technology Data Exchange (ETDEWEB)

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

    1996-12-31

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

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

    Energy Technology Data Exchange (ETDEWEB)

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

    1996-12-31

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

  17. Thrust stand for vertically oriented electric propulsion performance evaluation

    Energy Technology Data Exchange (ETDEWEB)

    Moeller, Trevor [University of Tennessee Space Institute, Tullahoma, Tennessee 37388 (United States); Polzin, Kurt A. [NASA, Marshall Space Flight Center, Huntsville, Alabama 35812 (United States)

    2010-11-15

    A variation of a hanging pendulum thrust stand capable of measuring the performance of an electric thruster operating in the vertical orientation is presented. The vertical orientation of the thruster dictates that the thruster must be horizontally offset from the pendulum pivot arm, necessitating the use of a counterweight system to provide a neutrally stable system. Motion of the pendulum arm is transferred through a balance mechanism to a secondary arm on which deflection is measured. A noncontact light-based transducer is used to measure displacement of the secondary beam. The members experience very little friction, rotating on twisting torsional pivots with oscillatory motion attenuated by a passive, eddy-current damper. Displacement is calibrated using an in situ thrust calibration system. Thermal management and self-leveling systems are incorporated to mitigate thermal and mechanical drifts. Gravitational force and torsional spring constants associated with flexure pivots provide restoring moments. An analysis of the design indicates that the thrust measurement range spans roughly four decades, with the stand capable of measuring thrust up to 12 N for a 200 kg thruster and up to approximately 800 mN for a 10 kg thruster. Data obtained from calibration tests performed using a 26.8 lbm simulated thruster indicated a resolution of 1 mN on 100 mN level thrusts, while those tests conducted on a 200 lbm thruster yielded a resolution of roughly 2.5 mN at thrust levels of 0.5 N and greater.

  18. Thrust Stand for Vertically Oriented Electric Propulsion Performance Evaluation

    Science.gov (United States)

    Moeller, Trevor; Polzin, Kurt A.

    2010-01-01

    A variation of a hanging pendulum thrust stand capable of measuring the performance of an electric thruster operating in the vertical orientation is presented. The vertical orientation of the thruster dictates that the thruster must be horizontally offset from the pendulum pivot arm, necessitating the use of a counterweight system to provide a neutrally-stable system. Motion of the pendulum arm is transferred through a balance mechanism to a secondary arm on which deflection is measured. A non-contact light-based transducer is used to measure displacement of the secondary beam. The members experience very little friction, rotating on twisting torsional pivots with oscillatory motion attenuated by a passive, eddy current damper. Displacement is calibrated using an in situ thrust calibration system. Thermal management and self-leveling systems are incorporated to mitigate thermal and mechanical drifts. Gravitational restoring force and torsional spring constants associated with flexure pivots provide restoring moments. An analysis of the design indicates that the thrust measurement range spans roughly four decades, with the stand capable of measuring thrust up to 12 N for a 200 kg thruster and up to approximately 800 mN for a 10 kg thruster. Data obtained from calibration tests performed using a 26.8 lbm simulated thruster indicated a resolution of 1 mN on 100 mN-level thrusts, while those tests conducted on 200 lbm thruster yielded a resolution of roughly 2.5 micro at thrust levels of 0.5 N and greater.

  19. Thrust stand for vertically oriented electric propulsion performance evaluation.

    Science.gov (United States)

    Moeller, Trevor; Polzin, Kurt A

    2010-11-01

    A variation of a hanging pendulum thrust stand capable of measuring the performance of an electric thruster operating in the vertical orientation is presented. The vertical orientation of the thruster dictates that the thruster must be horizontally offset from the pendulum pivot arm, necessitating the use of a counterweight system to provide a neutrally stable system. Motion of the pendulum arm is transferred through a balance mechanism to a secondary arm on which deflection is measured. A noncontact light-based transducer is used to measure displacement of the secondary beam. The members experience very little friction, rotating on twisting torsional pivots with oscillatory motion attenuated by a passive, eddy-current damper. Displacement is calibrated using an in situ thrust calibration system. Thermal management and self-leveling systems are incorporated to mitigate thermal and mechanical drifts. Gravitational force and torsional spring constants associated with flexure pivots provide restoring moments. An analysis of the design indicates that the thrust measurement range spans roughly four decades, with the stand capable of measuring thrust up to 12 N for a 200 kg thruster and up to approximately 800 mN for a 10 kg thruster. Data obtained from calibration tests performed using a 26.8 lbm simulated thruster indicated a resolution of 1 mN on 100 mN level thrusts, while those tests conducted on a 200 lbm thruster yielded a resolution of roughly 2.5 mN at thrust levels of 0.5 N and greater.

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

    Science.gov (United States)

    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

  1. Satellite propulsion spectral signature detection and analysis through Hall effect thruster plume and atmospheric modeling

    Science.gov (United States)

    Wheeler, Pamela; Cobb, Richard; Hartsfield, Carl; Prince, Benjamin

    2016-09-01

    Space Situational Awareness (SSA) is of utmost importance in today's congested and contested space environment. Satellites must perform orbital corrections for station keeping, devices like high efficiency electric propulsion systems such as a Hall effect thrusters (HETs) to accomplish this are on the rise. The health of this system is extremely important to ensure the satellite can maintain proper position and perform its intended mission. Electron temperature is a commonly used diagnostic to determine the efficiency of a hall thruster. Recent papers have coordinated near infrared (NIR) spectral measurements of emission lines in xenon and krypton to electron temperature measurements. Ground based observations of these spectral lines could allow the health of the thruster to be determined while the satellite is in operation. Another issue worth considering is the availability of SSA assets for ground-based observations. The current SSA architecture is limited and task saturated. If smaller telescopes, like those at universities, could successfully detect these signatures they could augment data collection for the SSA network. To facilitate this, precise atmospheric modeling must be used to pull out the signature. Within the atmosphere, the NIR has a higher transmission ratio and typical HET propellants are approximately 3x the intensity in the NIR versus the visible spectrum making it ideal for ground based observations. The proposed research will focus on developing a model to determine xenon and krypton signatures through the atmosphere and estimate the efficacy through ground-based observations. The model will take power modes, orbit geometries, and satellite altitudes into consideration and be correlated with lab and field observations.

  2. High Performance Plasma Channel Insulators for High Power Hall Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — NASA missions for planetary exploration require high power, long-life Hall thrusters. However, thruster power and lifetime are limited by the erosion of plasma...

  3. Micro Pulsed Inductive Thruster with Solid Fuel Option (uPIT_SF) Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The Micro Pulsed Inductive Thruster with Solid Fuel Option (5PIT_SF) is a high-precision impulse bit electromagnetic plasma micro-thruster. The 5PIT prototype is a...

  4. Lifetime Improvement of Large Scale Green Monopropellant Thrusters via Novel, Long-Life Catalysts Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop and life-test a flight-weight, 5N class green monopropellant thruster in Phase II. The most important feature that sets this thruster apart...

  5. High Performance Plasma Channel Insulators for High Power Hall Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — NASA missions for planetary exploration require high power, long-life Hall thrusters. However, thruster power and lifetime are limited by the erosion of plasma...

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

    Science.gov (United States)

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

    2016-07-01

    The 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

  7. Modeling of plasma in a hybrid electric propulsion for small satellites

    Science.gov (United States)

    Jugroot, Manish; Christou, Alex

    2016-09-01

    As space flight becomes more available and reliable, space-based technology is allowing for smaller and more cost-effective satellites to be produced. Working in large swarms, many small satellites can provide additional capabilities while reducing risk. These satellites require efficient, long term propulsion for manoeuvres, orbit maintenance and de-orbiting. The high exhaust velocity and propellant efficiency of electric propulsion makes it ideally suited for low thrust missions. The two dominant types of electric propulsion, namely ion thrusters and Hall thrusters, excel in different mission types. In this work, a novel electric hybrid propulsion design is modelled to enhance understanding of key phenomena and evaluate performance. Specifically, the modelled hybrid thruster seeks to overcome issues with existing Ion and Hall thruster designs. Scaling issues and optimization of the design will be discussed and will investigate a conceptual design of a hybrid spacecraft plasma engine.

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

    Science.gov (United States)

    2014-07-15

    characterize the spoke in a wide range of HETs, including both conventional and non-conventional designs (the H6 thruster, the NASA 173Mv1, the Busek BHT -600...near plume of the thruster[126]. Similarly, Liu [127, 128] also finds azimuthal oscillations in the BHT -200 and BHT -600 thrusters via high speed-imaging

  9. Study of Gas and Plasma Conditions in the High Isp VASIMR Thruster

    Science.gov (United States)

    Batishchev, O.; Molvig, K.

    2002-01-01

    -17, 1999, Torino, Italy, (1999) 99-102. [4] F.R.Chang Diaz et al., "An Overview of Current Research on the VASIMR Engine", DPP-2000, Bull. APS, vol.45, 7 (2000) 129. [5] J.P.Squire, "Recent Experimental Results in the VX-10 Device", DPP-2000, Bull. APS, vol.45, 7 (2000) 130. [6] O.Batishchev and K.Molvig, "Kinetic Simulation of the high Isp Plasma Thruster", JPC-36, AIAA-3754 technical paper, Huntsville (2000) 11p. [7] O.Batishchev and K.Molvig, "Study of the Operational Regimes of the VASIMR Helicon Plasma Source", DPP- 2000, Quebec City, Canada, Bull. APS, 45, 7 (2000) 130. [8] O.Batishchev and K.Molvig, "Study of Mixed Collisionality Gas Flow in the VASIMR Thruster", DFD-2000, DC, USA, Bull. APS, 45, 9 (2000) 169. [9] O.Batishchev and K.Molvig, "Kinetic Model of a Helicon Plasma Source for VASIMR", ACME-39, AIAA-0963 technical paper, ASME-39, Reno, 2001. [10] O.Batishchev and K.Molvig "Kinetic study of the VASIMR thruster operational regimes", JPC-37, AIAA-3501 technical paper, Salt-Lake City, 2001. [11] O.Batishchev and K.Molvig, "Numerical study of plasma production in the VASIMR thruster", IEPC-01-208 paper, 27 Int. Electric Prop. Conf., Pasadena CA, 15-19 October, -19p, 2001.

  10. Particle-in-cell simulation for different magnetic mirror effects on the plasma distribution in a cusped field thruster

    Science.gov (United States)

    Liu, Hui; Chen, Peng-Bo; Zhao, Yin-Jian; Yu, Da-Ren

    2015-08-01

    Magnetic mirror used as an efficient tool to confine plasma has been widely adopted in many different areas especially in recent cusped field thrusters. In order to check the influence of magnetic mirror effect on the plasma distribution in a cusped field thruster, three different radii of the discharge channel (6 mm, 4 mm, and 2 mm) in a cusped field thruster are investigated by using Particle-in-Cell Plus Monte Carlo (PIC-MCC) simulated method, under the condition of a fixed axial length of the discharge channel and the same operating parameters. It is found that magnetic cusps inside the small radius discharge channel cannot confine electrons very well. Thus, the electric field is hard to establish. With the reduction of the discharge channel’s diameter, more electrons will escape from cusps to the centerline area near the anode due to a lower magnetic mirror ratio. Meanwhile, the leak width of the cusped magnetic field will increase at the cusp. By increasing the magnetic field strength in a small radius model of a cusped field thruster, the negative effect caused by the weak magnetic mirror effect can be partially compensated. Therefore, according to engineering design, the increase of magnetic field strength can contribute to obtaining a good performance, when the radial distance between the magnets and the inner surface of the discharge channel is relatively big. Project supported by the National Natural Science Foundation of China (Grant No. 51006028) and the Foundation for Innovative Research Groups of the National Natural Science Foundation of China (Grant No. 51121004).

  11. Primary electric propulsion for future space missions

    Science.gov (United States)

    Byers, D. C.; Terdan, F. F.; Myers, I. T.

    1979-01-01

    The paper presents data and a methodology to allow preliminary definition of electric propulsion systems. The elements comprising the thrust system are described parametrically. As an example, thruster performance is presented as a function of specific impulse and propellant type. Two power management and control (PMAC) approaches are considered to illustrate the use of the methodology. Power source characteristics are disregarded in the system description. One PMAC concept assumes a dc power input to the thrust system and all thruster power conditioned in a conventional manner; the other PMAC approach assumes an ac power source.

  12. Continuous Wheel Momentum Dumping Using Magnetic Torquers and Thrusters

    Science.gov (United States)

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

    1996-12-01

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

  13. The electrodeless Lorentz force (ELF) thruster experimental facility

    Science.gov (United States)

    Weber, T. E.; Slough, J. T.; Kirtley, D.

    2012-11-01

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

  14. 2-D Magnetohydrodynamic Modeling of A Pulsed Plasma Thruster

    Science.gov (United States)

    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.

  15. Experimental characterization of radio frequency microthermal thruster performance

    Science.gov (United States)

    Williams, Shae E.

    Microsatellite (cold gas thrusters. Design constraints rule out much of traditional propulsion, requiring new and nonobvious technologies to advance the state of the art and enable longer and more flexible missions. The radio frequency microthermal thruster is shown to be worth thorough study for this application. A basic analytical model is constructed to look at expected performance, and the theory behind that model is explained. Calibration and the challenges in working with extremely low forces and displacements are also examined. The results of extensive testing on this thruster type are presented. Important trends are confirmed and validated, such as a linearity of specific impulse with power, and consistent nonlinearities with frequency and mass flow rate. Additionally, tests indicate a nonlinear relationship between applied frequency and thruster internal geometry that can more than triple the heating occurring in the thruster. Further tests focus on this relationship, and find more information about how these parameters couple are found to be primarily due to induced inefficiencies in stochastic heating and the inability of a vibrating voltage sheath to transfer energy into the flow. Additionally, first steps towards optimizing a design for performance are taken, such as analyzing the effect of adding a converging/diverging nozzle and finding an optimal length of inner electrode to be exposed to plasma. Overall, specific impulses of up to 85 seconds are found with argon as the propellant, doubling cold gas specific impulse, and an error on specific impulse is calculated to be less than 3% in either direction. These results after only slight efforts at design optimization indicate much more improvement is possible with this technology that would make an RF microthermal thruster viable as a commercial product.

  16. Enhancing space transportation: The NASA program to develop electric propulsion

    Science.gov (United States)

    Bennett, Gary L.; Watkins, Marcus A.; Byers, David C.; Barnett, John W.

    1990-01-01

    The NASA Office of Aeronautics, Exploration, and Technology (OAET) supports a research and technology (R and T) program in electric propulsion to provide the basis for increased performance and life of electric thruster systems which can have a major impact on space system performance, including orbital transfer, stationkeeping, and planetary exploration. The program is oriented toward providing high-performance options that will be applicable to a broad range of near-term and far-term missions and vehicles. The program, which is being conducted through the Jet Propulsion Laboratory (JPL) and Lewis Research Center (LeRC) includes research on resistojet, arcjets, ion engines, magnetoplasmadynamic (MPD) thrusters, and electrodeless thrusters. Planning is also under way for nuclear electric propulsion (NEP) as part of the Space Exploration Initiative (SEI).

  17. X-ray emission from a nanosecond-pulse discharge in an inhomogeneous electric field at atmospheric pressure

    Energy Technology Data Exchange (ETDEWEB)

    Zhang Cheng; Shao Tao; Ren Chengyan; Zhang Dongdong [Institute of Electrical Engineering, Chinese Academy of Sciences, Beijing 100190 (China); Key Laboratory of Power Electronics and Electric Drive, Chinese Academy of Sciences, Beijing 100190 (China); Tarasenko, Victor; Kostyrya, Igor D. [Institute of High Current Electronics, Russian Academy of Science, Tomsk 634055 (Russian Federation); Ma Hao [Institute of Electrical Engineering, Chinese Academy of Sciences, Beijing 100190 (China); Yan Ping [Institute of Electrical Engineering, Chinese Academy of Sciences, Beijing 100190 (China); Key Laboratory of Power Electronics and Electric Drive, Chinese Academy of Sciences, Beijing 100190 (China); State Key Laboratory of Electrical Insulation and Power Equipment, Xi' an Jiaotong University, Xi' an 710049 (China)

    2012-12-15

    This paper describes experimental studies of the dependence of the X-ray intensity on the anode material in nanosecond high-voltage discharges. The discharges were generated by two nanosecond-pulse generators in atmospheric air with a highly inhomogeneous electric field by a tube-plate gap. The output pulse of the first generator (repetitive pulse generator) has a rise time of about 15 ns and a full width at half maximum of 30-40 ns. The output of the second generator (single pulse generator) has a rise time of about 0.3 ns and a full width at half maximum of 1 ns. The electrical characteristics and the X-ray emission of nanosecond-pulse discharge in atmospheric air are studied by the measurement of voltage-current waveforms, discharge images, X-ray count and dose. Our experimental results showed that the anode material rarely affects electrical characteristics, but it can significantly affect the X-ray density. Comparing the density of X-rays, it was shown that the highest x-rays density occurred in the diffuse discharge in repetitive pulse mode, then the spark discharge with a small air gap, and then the corona discharge with a large air gap, in which the X-ray density was the lowest. Therefore, it could be confirmed that the bremsstrahlung at the anode contributes to the X-ray emission from nanosecond-pulse discharges.

  18. Plasma Thruster Development: Magnetoplasmadynamic Propulsion, Status and Basic Problems.

    Science.gov (United States)

    1986-02-01

    Closed Drift Hall-Ion Thruster Flown on the Russian Satellite Meteor I, 1971, from Reference 13 12 4 Flat Coil Induction Thruster Schematic from...the Russian Satellite Meteor 1, 1971. from Ref. 1-3. 13 COIL Br PLASMA SWITCH0 0 FZ jeBr 0 CAPACITOR 0 Fig.- 4:Fa olInuto huse ceai fromRef-22 40 14 is...minute crater (on the order of 10- 4 cm diameter). High pressures, on the order of 100 bar, and vaporization rates in these craters have been

  19. Physics and Dynamics of Current Sheets in Pulsed Plasma Thrusters

    Science.gov (United States)

    2007-11-02

    pulsed plasma thruster. A simple experiment would involve measuring the impulse bit of a coaxial gas-fed pulsed plasma thruster operated in both positive...Princeton, NJ, 2002. [2] J. Marshal. Performance of a hydromagnetic plasma gun . The Physics of Fluids, 3(1):134–135, January-February 1960. [3] R.G. Jahn...Jahn and K.E. Clark. A large dielecteic vacuum facility. AIAA Jour- nal, 1966. [16] L.C. Burkhardt and R.H. Lovberg. Current sheet in a coaxial plasma

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

    Science.gov (United States)

    Cannat, F.; Lafleur, T.; Jarrige, J.; Chabert, P.; Elias, P.-Q.; Packan, D.

    2015-05-01

    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.

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2015-05-15

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

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

    Directory of Open Access Journals (Sweden)

    Zool H. Ismail

    2014-01-01

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

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    Pencil, Eirc S.; Benson, Scott W.

    2008-01-01

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

  5. Thruster direction controlling of assembled spacecraft based on gimbal suspension

    Institute of Scientific and Technical Information of China (English)

    Hongliang Xu; Hai Huang

    2016-01-01

    The attitude control system design and its control effect are affected considerably by the mass-property pa-rameters of the spacecraft. In the mission of on-orbit servicing, as fuel is expended, or the payloads are added or removed, the center of mass wil be changed in certain axe; conse-quently, some thrusters' directions are deviated from the center of mass (CM) in certain plane. The CM of assembled spacecraft estimation and thruster direction control are studied. Firstly, the attitude dynamics of the assembled spacecraft is established based on the Newton-Euler method. Secondly, the estimation can be identified by the least recursive squares algorithm. Then, a scheme to control the thrusters’ directions is proposed. By using the gimbal instaled at the end of the boom, the angle of the thruster is controled by driving the gimbal; therefore, thrusters can be directed to the CM again. Finaly, numerical simulations are used to verify this scheme. Results of the numerical simulations clearly show that this control scheme is rational and feasible.

  6. Mode Transitions in Magnetically Shielded Hall Effect Thrusters

    Science.gov (United States)

    Sekerak, Michael J.; Longmier, Benjamin W.; Gallimore, Alec D.; Huang, Wensheng; Kamhawi, Hani; Hofer, Richard R.; Jorns, Benjamin A.; Polk, James E.

    2014-01-01

    A mode transition study is conducted in magnetically shielded thrusters where the magnetic field magnitude is varied to induce mode transitions. Three different oscillatory modes are identified with the 20-kW NASA-300MS-2 and the 6-kW H6MS: Mode 1) global mode similar to unshielded thrusters at low magnetic fields, Mode 2) cathode oscillations at nominal magnetic fields, and Mode 3) combined spoke, cathode and breathing mode oscillations at high magnetic fields. Mode 1 exhibits large amplitude, low frequency (1-10 kHz), breathing mode type oscillations where discharge current mean value and oscillation amplitude peak. The mean discharge current is minimized while thrust-to-power and anode efficiency are maximized in Mode 2, where higher frequency (50-90 kHz), low amplitude, cathode oscillations dominate. Thrust is maximized in Mode 3 and decreases by 5-6% with decreasing magnetic field strength. The presence or absence of spokes and strong cathode oscillations do not affect each other or discharge current. Similar to unshielded thrusters, mode transitions and plasma oscillations affect magnetically shielded thruster performance and should be characterized during system development.

  7. STS-39: OMS Pod Thruster Removal/Replace

    Science.gov (United States)

    1991-01-01

    Shown is the removal and replacement of the Discovery's orbital maneuvering systems (OMS) pod thruster. The OMS engine will be used to propel Discovery north, off of its previous orbital groundtrack, without changing the spacecraft's altitude. A burn with this lateral effect is known as "out-of-plane."

  8. Thermal stability of the krypton Hall effect thruster

    Directory of Open Access Journals (Sweden)

    Szelecka Agnieszka

    2017-03-01

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

  9. Simulations of a Plasma Thruster Utilizing the FRC Configuration

    Energy Technology Data Exchange (ETDEWEB)

    Rognlien, T. D. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States); Cohen, B. I. [Lawrence Livermore National Lab. (LLNL), Livermore, CA (United States)

    2016-10-10

    This report describes work performed by LLNL to model the behavior and performance of a reverse-field configuration (FRC) type of plasma device as a plasma thruster as summarized by Razin et al. [1], which also describes the MNX device at PPPL used to study this concept.

  10. Theory for the anomalous electron transport in Hall-effect thrusters

    Science.gov (United States)

    Lafleur, Trevor; Baalrud, Scott; Chabert, Pascal

    2016-09-01

    Using insights from particle-in-cell (PIC) simulations, we develop a kinetic theory to explain the anomalous cross-field electron transport in Hall-effect thrusters (HETs). The large axial electric field in the acceleration region of HETs, together with the radially applied magnetic field, causes electrons to drift in the azimuthal direction with a very high velocity. This drives an electron cyclotron instability that produces large amplitude oscillations in the plasma density and azimuthal electric field, and which is convected downstream due to the large axial ion drift velocity. The frequency and wavelength of the instability are of the order of 5 MHz and 1 mm respectively, while the electric field amplitude can be of a similar magnitude to axial electric field itself. The instability leads to enhanced electron scattering many orders of magnitude higher than that from standard electron-neutral or electron-ion Coulomb collisions, and gives electron mobilities in good agreement with experiment. Since the instability is a strong function of almost all plasma properties, the mobility cannot in general be fitted with simple 1/B or 1/B2 scaling laws, and changes to the secondary electron emission coefficient of the HET channel walls are expected to play a role in the evolution of the instability. This work received financial support from a CNES postdoctoral research award.

  11. Engineering Model Propellant Feed System Development for an Iodine Hall Thruster Demonstration Mission

    Science.gov (United States)

    Polzin, Kurt A.

    2016-01-01

    CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cu cm and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high (Delta)v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ?Isp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (< 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum

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

    Science.gov (United States)

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

    2010-01-01

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

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

    Science.gov (United States)

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

    2015-01-01

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

  14. Interaction of Repetitively Pulsed High Energy Laser Radiation With Matter

    Science.gov (United States)

    Hugenschmidt, Manfred

    1986-10-01

    The paper is concerned with laser target interaction processes involving new methods of improving the overall energy balance. As expected theoretically, this can be achieved with high repetition rate pulsed lasers even for initially highly reflecting materials, such as metals. Experiments were performed by using a pulsed CO2 laser at mean powers up to 2 kW and repetition rates up to 100 Hz. The rates of temperature rise of aluminium for example were thereby increased by lore than a factor of 3 as compared to cw-radiation of comparable power density. Similar improvements were found for the overall absorptivities that were increased by this method by more than an order of magnitude.

  15. Characterization of advanced electric propulsion systems

    Science.gov (United States)

    Ray, P. K.

    1982-01-01

    Characteristics of several advanced electric propulsion systems are evaluated and compared. The propulsion systems studied are mass driver, rail gun, MPD thruster, hydrogen free radical thruster and mercury electron bombardment ion engine. These are characterized by specific impulse, overall efficiency, input power, average thrust, power to average thrust ratio and average thrust to dry weight ratio. Several important physical characteristics such as dry system mass, accelerator length, bore size and current pulse requirement are also evaluated in appropriate cases. Only the ion engine can operate at a specific impulse beyond 2000 sec. Rail gun, MPD thruster and free radical thruster are currently characterized by low efficiencies. Mass drivers have the best performance characteristics in terms of overall efficiency, power to average thrust ratio and average thrust to dry weight ratio. But, they can only operate at low specific impulses due to large power requirements and are extremely long due to limitations of driving current. Mercury ion engines have the next best performance characteristics while operating at higher specific impulses. It is concluded that, overall, ion engines have somewhat better characteristics as compared to the other electric propulsion systems.

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

    Science.gov (United States)

    Mirtich, M. J.

    1976-01-01

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

  17. Thermal-environment testing of a 30-cm engineering model thruster

    Science.gov (United States)

    Mirtich, M. J.

    1976-01-01

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

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

    Science.gov (United States)

    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.

  19. Measurement of axial neutral density profiles in a microwave discharge ion thruster by laser absorption spectroscopy with optical fiber probes.

    Science.gov (United States)

    Tsukizaki, Ryudo; Koizumi, Hiroyuki; Nishiyama, Kazutaka; Kuninaka, Hitoshi

    2011-12-01

    In order to reveal the physical processes taking place within the "μ10" microwave discharge ion thruster, internal plasma diagnosis is indispensable. However, the ability of metallic probes to access microwave plasmas biased at a high voltage is limited from the standpoints of the disturbance created in the electric field and electrical isolation. In this study, the axial density profiles of excited neutral xenon were successfully measured under ion beam acceleration by using a novel laser absorption spectroscopy system. The target of the measurement was metastable Xe I 5p(5)((2)P(0) (3/2))6s[3/2](0) (2) which absorbed a wavelength of 823.16 nm. Signals from laser absorption spectroscopy that swept a single-mode optical fiber probe along the line of sight were differentiated and converted into axial number densities of the metastable neutral particles in the plasma source. These measurements revealed a 10(18) m(-3) order of metastable neutral particles situated in the waveguide, which caused two different modes during the operation of the μ10 thruster. This paper reports a novel spectroscopic measurement system with axial resolution for microwave plasma sources utilizing optical fiber probes.

  20. Development, Vibration, and Thermal Characterization of a Steady Operating Pulsed Power System for FRC Thrusters

    Science.gov (United States)

    2015-04-01

    Field (RMF) to produce large plasma currents inside a conical thruster creating a field-reversed configuration (FRC) plasmoid that is magnetically...in turn charges a high-Q capacitor. Connected in series with the thruster antenna, the resonant RLC circuit oscillates at high frequency with a...Field (RMF) to produce large plasma currents inside a conical thruster creating a field-reversed configuration (FRC) plasmoid that is magnetically

  1. Colloid Thruster for Attitude Control Systems (ACS) and Tip-off Control Applications Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop and deliver a complete engineering model colloid thruster system, capable of thrust levels and lifetimes required for spacecraft...

  2. Low-Cost High-Performance Hall Thruster Support System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Colorado Power Electronics (CPE) has built an innovative modular power processing unit (PPU) for Hall Thrusters, including discharge, magnet, heater and keeper...

  3. Plasma Characterization of Hall Thruster with Active and Passive Segmented Electrodes

    Energy Technology Data Exchange (ETDEWEB)

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

    2002-09-04

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

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

    Science.gov (United States)

    Jorns, Benjamin A.; Hofery, Richard R.

    2013-01-01

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

  5. Design of a Laboratory Hall Thruster with Magnetically Shielded Channel Walls, Phase I: Numerical Simulations

    Science.gov (United States)

    Mikellides, Ioannis G.; Katz, Ira; Hofer, Richard R.

    2011-01-01

    In a proof-of-principle effort to demonstrate the feasibility of magnetically shielded (MS) Hall thrusters, an existing laboratory thruster has been modified with the guidance of physics-based numerical simulation. When operated at a discharge power of 6-kilowatts the modified thruster has been designed to reduce the total energy and flux of ions to the channel insulators by greater than 1 and greater than 3 orders of magnitude, respectively. The erosion rates in this MS thruster configuration are predicted to be at least 2-4 orders of magnitude lower than those in the baseline (BL) configuration. At such rates no detectable erosion is expected to occur.

  6. MOA—The Magnetic Field Amplified Thruster, a Novel Concept for a Pulsed Plasma Accelerator

    Science.gov (United States)

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

    2008-01-01

    More than 60 years after the later Nobel laureate Hannes Alfvén 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 Alfvén 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&D activities at QASAR (www.qasar.at), the company in Vienna, which has been set up to further develop and test the Alfvén wave technology and its applications.

  7. Studies of anode sheath phenomena in a Hall-effect thruster discharge

    Science.gov (United States)

    Dorf, Leonid

    2005-10-01

    Crossed electric and magnetic fields devices (plasma thrusters, magnetrons, coaxial plasma guns, plasma opening switches, etc.) are routinely used for plasma production and in other applications. Despite these numerous applications, the fundamental anode sheath phenomena in many of these devices have received surprisingly little experimental scrutiny. We chose a Hall-effect thruster (HT) discharge for our study of the anode sheath. It has been typically assumed in most fluid models of an HT that its steady-state operation requires the presence of a negative anode fall (electron-repelling anode sheath). Such anode fall behavior, opposite to that in typical glow discharges or hollow-anode plasma sources, is the result of a relatively high degree of ionization in HTs, achieved by applying a radial magnetic field transverse to the direction of the discharge current. Our data from non-perturbing probe measurements showed for the first time that the anode fall in HTs can be either negative or positive (electron-attracting anode sheath), depending on conditions at the anode surface. The path for current closure to the anode turns out to be quite subtle in HTs. This path determines the mechanism of the anode fall formation. In varying the magnetic field topology in the channel from a more uniform to a cusp-like one, we uncover intriguing results. For cusp configurations, in which the radial magnetic field changes polarity somewhere along the channel, the anode fall is positive, whereas it is negative for a more uniform field. This polarity difference could be attributed to the decreased electron mobility across the magnetic field in the cusp-like configuration. Our theoretical modeling of the anode sheath correlates well with the experimental results in describing how the magnitude of the sheath varies with the discharge voltage and mass flow rate.

  8. Geometrical characterization and performance optimization of monopropellant thruster injector

    Directory of Open Access Journals (Sweden)

    T.R. Nada

    2012-12-01

    Full Text Available The function of the injector in a monopropellant thruster is to atomize the liquid hydrazine and to distribute it over the catalyst bed as uniformly as possible. A second objective is to place the maximum amount of catalyst in contact with the propellant in as short time as possible to minimize the starting transient time. Coverage by the spray is controlled mainly by cone angle and diameter of the catalyst bed, while atomization quality is measured by the Sauter Mean Diameter, SMD. These parameters are evaluated using empirical formulae. In this paper, two main types of injectors are investigated; plain orifice and full cone pressure swirl injectors. The performance of these two types is examined for use with blow down monopropellant propulsion system. A comprehensive characterization is given and design charts are introduced to facilitate optimizing the performance of the injector. Full-cone injector is a more suitable choice for monopropellant thruster and it might be available commercially.

  9. Convective heat flux in a laser-heated thruster

    Science.gov (United States)

    Wu, P. K. S.

    1978-01-01

    An analysis is performed to estimate the convective heating to the wall in a laser-heated thruster on the basis of a solution of the laminar boundary-layer equations with variable transport properties. A local similiarity approximation is used, and it is assumed that the gas phase is in equilibrium. For the thruster described by Wu (1976), the temperature and pressure distributions along the nozzle are obtained from the core calculation. The similarity solutions and heat flux are obtained from the freestream conditions of the boundary layer, in order to determine if it is necessary to couple the boundary losses directly to the core calculation. In addition, the effects of mass injection on the convective heat transfer across the boundary layer with large density-viscosity product gradient are examined.

  10. Controllability of an underactuated spacecraft with one thruster under disturbance

    Institute of Scientific and Technical Information of China (English)

    Dong-Xia Wang; Ying-Hong Jia; Lei Jin; Hai-Chao Gui; Shi-Jie Xu

    2012-01-01

    For an underactuated spacecraft using only one thruster,the attitude controllability with respect to the orbit frame is studied in the presence of periodical oscillation disturbance,which provides a preconditional guide on designing control law for underactuated attitude control system.Firstly,attitude dynamic model was established for an underactuated spacecraft,and attitude motion was described using the special orthogonal group (SO (3)).Secondly,Liouville theorem was used to confirm that the flow generated by the drift vector of the underactuated attitude control system is volume-preserving.Furthermore,according to Poincaré's recurrence theorem,we draw conclusions that this drift field is weakly positively poisson stable (WPPS).Thirdly,the sufficient and necessary condition of controllability was obtained on the basis of lie algebra rank condition (LARC).Finally,the controllable conditions were analyzed and simulated in different cases of inertia matrix with the installed position of thruster.

  11. Thrusters Pairing Guidelines for Trajectory Corrections of Projectiles

    Science.gov (United States)

    2011-01-12

    Gill, J., “Experimental Investigation of Super- and Hypersonic Jet Interaction on Missile Configurations,” Journal of Spacecraft and Rockets, Vol. 35...Thrusters Pairing Guidelines for Trajectory Corrections of Projectiles Daniel Corriveau∗ Canadian Department of National Defence , Quebec City, Quebec...course correction process for a 30-mm fin-stabilized air- defense projectile and a standard 105-mm spin-stabilized artillery shell are presented

  12. Investigation of Hall Effect Thruster Channel Wall Erosion Mechanisms

    Science.gov (United States)

    2016-08-01

    mentorship, humor, and amazing barbecue parties. I would also like to thank my thesis committee, Dr. Yim, Prof Simon, and Dr. Ready for taking the...indicate that BN is depleted relative to silica in the highly eroded region of the thruster. This surprising result mirrors that obtained by Garnier...predict the decrease in BN in the HE region. Grain ejection provides a plausible mechanism that could explain this surprising observation. 3.6. Summary

  13. Hall Effect Thruster Plume Contamination and Erosion Study

    Science.gov (United States)

    Jaworske, Donald A.

    2000-01-01

    The objective of the Hall effect thruster plume contamination and erosion study was to evaluate the impact of a xenon ion plume on various samples placed in the vicinity of a Hall effect thruster for a continuous 100 hour exposure. NASA Glenn Research Center was responsible for the pre- and post-test evaluation of three sample types placed around the thruster: solar cell cover glass, RTV silicone, and Kapton(R). Mass and profilometer), were used to identify the degree of deposition and/or erosion on the solar cell cover glass, RTV silicone, and Kapton@ samples. Transmittance, reflectance, solar absorptance, and room temperature emittance were used to identify the degree of performance degradation of the solar cell cover glass samples alone. Auger spectroscopy was used to identify the chemical constituents found on the surface of the exposed solar cell cover glass samples. Chemical analysis indicated some boron nitride contamination on the samples, from boron nitride insulators used in the body of the thruster. However, erosion outweighted contamination. All samples exhibited some degree of erosion. with the most erosion occurring near the centerline of the plume and the least occurring at the +/- 90 deg positions. For the solar cell cover glass samples, erosion progressed through the antireflective coating and into the microsheet glass itself. Erosion occurred in the solar cell cover glass, RTV silicone and Kapton(R) at different rates. All optical properties changed with the degree of erosion, with solar absorptance and room temperature emittance increasing with erosion. The transmittance of some samples decreased while the reflectance of some samples increased and others decreased. All results are consistent with an energetic plume of xenon ions serving as a source for erosion.

  14. Vacuum arc plasma thrusters with inductive energy storage driver

    Science.gov (United States)

    Krishnan, Mahadevan (Inventor)

    2009-01-01

    A plasma thruster with a cylindrical inner and cylindrical outer electrode generates plasma particles from the application of energy stored in an inductor to a surface suitable for the formation of a plasma and expansion of plasma particles. The plasma production results in the generation of charged particles suitable for generating a reaction force, and the charged particles are guided by a magnetic field produced by the same inductor used to store the energy used to form the plasma.

  15. Iodine Plasma Species Measurements in a Hall Effect Thruster Plume

    Science.gov (United States)

    2013-05-01

    60 90 0 2 4 6 8 Current (mA/cm^2) A n g l e ( d e g ) Xenon Iodine 500 V, 2 A, I2 Presented at 2012 JPC 33 Distribution A: Approved for public...Over 1 hour of operation on iodine – Additional 1/2 hour with thruster flowing Xe – Current up to ~50 A into anode Presented at 2012 JPC

  16. Estimating Thruster Impulses From IMU and Doppler Data

    Science.gov (United States)

    Lisano, Michael E.; Kruizinga, Gerhard L.

    2009-01-01

    A computer program implements a thrust impulse measurement (TIM) filter, which processes data on changes in velocity and attitude of a spacecraft to estimate the small impulsive forces and torques exerted by the thrusters of the spacecraft reaction control system (RCS). The velocity-change data are obtained from line-of-sight-velocity data from Doppler measurements made from the Earth. The attitude-change data are the telemetered from an inertial measurement unit (IMU) aboard the spacecraft. The TIM filter estimates the threeaxis thrust vector for each RCS thruster, thereby enabling reduction of cumulative navigation error attributable to inaccurate prediction of thrust vectors. The filter has been augmented with a simple mathematical model to compensate for large temperature fluctuations in the spacecraft thruster catalyst bed in order to estimate thrust more accurately at deadbanding cold-firing levels. Also, rigorous consider-covariance estimation is applied in the TIM to account for the expected uncertainty in the moment of inertia and the location of the center of gravity of the spacecraft. The TIM filter was built with, and depends upon, a sigma-point consider-filter algorithm implemented in a Python-language computer program.

  17. Two-dimensional model of stationary plasma thruster

    Energy Technology Data Exchange (ETDEWEB)

    Pitchford, L.C.; Boeuf, J.P. [Universite Paul Sabatier, Toulouse (France)

    1995-12-31

    A stationary plasma thruster, SPT, (also called closed-drift thrusters or Hall thrusters) is an electromagnetic propulseur design which has been developed over the past thirty years in the former USSR. SPT`s are small devices with a thrust greater than 1000 s{sup -1}, and a lifetime of several 1000 hours or more. These properties make the SPT of interest for applications such as satellite station-keeping or orbit transfer. The geometry of the SPT is shown; it consists of a hollow, cylindrical dielectric (typically of several centimeters length and diameter) with a central dielectric rod. A voltage on the order of several 100`s of V is applied at the anode (at one end of the cylinder). The cathode is an externally powered hollow cathode or a hot filament positioned slightly past the exit of the dielectric cylinder. Gas, typically xenon, flows in from around the anode and is ionized by the electrons which are emitted from the cathode. A magnetic field is applied which is mainly in the radial direction. The magnetic strength is such that the electrons tend to be trapped along the magnetic field lines, but the ion trajectories are not significantly influenced by the magnetic field. For these conditions, the current at the anode is several amps. At the exit plane, the xenon is almost fully ionized. The ion flux at the exhaust provides the thrust.

  18. A novel laser ablation plasma thruster with electromagnetic acceleration

    Science.gov (United States)

    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.

  19. Experimental studies of anode sheath phenomena in a hall thruster.

    Energy Technology Data Exchange (ETDEWEB)

    Dorf, L. A. (Leonid A.); Fisch, N. J.; Raitses, Yevgeny F.

    2004-01-01

    Both electron-repelling (negative anode fall) and electron-attracting (positive anode fall) anode sheaths in a Hall thruster were identified experimentally by performing accurate, non-disturbing near-anode measurements with biased and emissive probes. An interesting new phenomenon revealed by the probe measurements is that the anode fall changes from positive to negative upon removal of the dielectric coating, which appears on the anode surface during the course of Hall thruster operation. Probe measurements in a Hall thruster with three different magnetic field configurations show that an anode fall at the clean anode is a function of the radial magnetic field profile inside the channel. A positive anode fall formation mechanism suggested in this work is that: (1) when the anode front surface is coated with dielectric, a discharge current closes to the anode at the surfaces that remain conductive, (2) a total thermal electron current toward the conductive area is significantly smaller than the discharge current, therefore an additional electron flux needs to be attracted toward the conductive surfaces by the electronattracting sheath that appears at these surfaces.

  20. Oral stereognostic ability among tongue thrusters with interdental lisp, tongue thrusters without interdental lisp and normal children.

    Science.gov (United States)

    Colletti, E A; Geffner, D; Schlanger, P

    1976-02-01

    30 children, i.e., 10 children per group, 8 yr. of age, were given an oral stereognostic test. This test of 10 geometric forms varying in shape were developed by NIDR. 47 stimuli pairs were used and 10 pairs were repeated to measure test reliability. Subjects were blindfolded and asked to respond whether Items 1 and 2, presented consecutively, were the same or different. Results indicated that both groups of tongue thrusters with and without interdental lisp scored significantly more poorly than did normal children (t = 4.68, P less than .001; t = 5.00, P less than .001), respectively. There were no significant differences, however, between tongue thrusters with and without interdental lisp (t = .33, P greater than .05). Observations indicated that normal children used the tongue tip more frequently and accurately when discriminating the geometric forms than did the other groups.

  1. Langmuir Probe Measurements Within the Discharge Channel of the 20-kW NASA-300M and NASA-300MS Hall Thrusters

    Science.gov (United States)

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

    2013-01-01

    NASA is presently developing a high-power, high-efficiency, long-lifetime Hall thruster for the Solar Electric Propulsion Technology Demonstration Mission. In support of this task, studies have been performed on the 20-kW NASA-300M Hall thruster to aid in the overall design process. The ability to incorporate magnetic shielding into a high-power Hall thruster was also investigated with the NASA- 300MS, a modified version of the NASA-300M. The inclusion of magnetic shielding would allow the thruster to push existing state-of-the-art technology in regards to service lifetime, one of the goals of the Technology Demonstration Mission. Langmuir probe measurements were taken within the discharge channels of both thrusters in order to characterize differences at higher power levels, as well as validate ongoing modeling efforts using the axisymmetric code Hall2De. Flush-mounted Langmuir probes were also used within the channel of the NASA-300MS to verify that magnetic shielding was successfully applied. Measurements taken from 300 V, 10 kW to 600 V, 20 kW have shown plasma potentials near anode potential and electron temperatures of 4 to 12 eV at the walls near the thruster exit plane of the NASA-300MS, verifying magnetic shielding and validating the design process at this power level. Channel centerline measurements on the NASA-300M from 300 V, 10 kW to 500 V, 20 kW show the electron temperature peak at approximately 0.1 to 0.2 channel lengths upstream of the exit plane, with magnitudes increasing with discharge voltage. The acceleration profiles appear to be centered about the exit plane with a width of approximately 0.3 to 0.4 channel lengths. Channel centerline measurements on the NASA-300MS were found to be more challenging due to additional probe heating. Ionization and acceleration zones appeared to move downstream on the NASA-300MS compared to the NASA-300M, as expected based on the shift in peak radial magnetic field. Additional measurements or alternative

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

    Science.gov (United States)

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

    2009-04-01

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

  3. Influence of Tailored Applied Magnetic Fields on High-Power MPD Thruster Current Transport and Onset-Related Phenomena

    Science.gov (United States)

    Moeller, Robert C.; Polk, James E.

    2013-01-01

    This work investigated the effects of tailored, externally-applied magnetic fields on current transport and near-anode processes in the plasma discharge of a magnetoplasmadynamic thruster (MPDT). Electrical and plasma diagnostics were used to investigate how localized applied magnetic fields could mitigate the effects of the "onset" phenomena, including large-amplitude terminal voltage fluctuations and high anode fall voltages associated with unstable operation and anode erosion. An MPDT with a multi-channel hollow cathode was developed and tested with quasi-steady pulses of 1 millisecond duration at power levels of 36 kilowatts (20 volts, 1800 amperes) to 3.3 milliwatts (255 volts, 13.1 kiloamperes) with argon propellant in three different magnetic configurations: self-field, applied B field tangential to the anode lip near the exit plane, and applied cusp B field. The current pattern and current densities redistributed to follow the applied poloidal magnetic field lines, which created increased conduction paths to the anode. Also, the anode fall voltage was substantially reduced with both applied B field topologies over a large range of currents. For example, at 10.7 kiloamperes, the cusp applied magnetic field decreased anode fall voltages from 45-83 volts down to 15 volts or lower along much of the anode. The amplitude and frequency of the voltage fluctuations were also reduced over a broad range of currents with the applied fields. E.g., the standard deviations of the fluctuations were lowered by 37-49 percent at 8-9 kiloamperes. In addition, decreases in the mean terminal voltages as large as 31 percent were measured with the applied magnetic fields. These effects are shown to be associated with the increased current conduction along the applied magnetic field lines in the near-anode region. These results also suggest a reduction in frequency and intensity of current-concentrating filaments and anode spots, which contribute to erosion. Overall, both applied

  4. Influence of Tailored Applied Magnetic Fields on High-Power MPD Thruster Current Transport and Onset-Related Phenomena

    Science.gov (United States)

    Moeller, Robert C.; Polk, James E.

    2013-01-01

    This work investigated the effects of tailored, externally-applied magnetic fields on current transport and near-anode processes in the plasma discharge of a magnetoplasmadynamic thruster (MPDT). Electrical and plasma diagnostics were used to investigate how localized applied magnetic fields could mitigate the effects of the "onset" phenomena, including large-amplitude terminal voltage fluctuations and high anode fall voltages associated with unstable operation and anode erosion. An MPDT with a multi-channel hollow cathode was developed and tested with quasi-steady pulses of 1 millisecond duration at power levels of 36 kilowatts (20 volts, 1800 amperes) to 3.3 milliwatts (255 volts, 13.1 kiloamperes) with argon propellant in three different magnetic configurations: self-field, applied B field tangential to the anode lip near the exit plane, and applied cusp B field. The current pattern and current densities redistributed to follow the applied poloidal magnetic field lines, which created increased conduction paths to the anode. Also, the anode fall voltage was substantially reduced with both applied B field topologies over a large range of currents. For example, at 10.7 kiloamperes, the cusp applied magnetic field decreased anode fall voltages from 45-83 volts down to 15 volts or lower along much of the anode. The amplitude and frequency of the voltage fluctuations were also reduced over a broad range of currents with the applied fields. E.g., the standard deviations of the fluctuations were lowered by 37-49 percent at 8-9 kiloamperes. In addition, decreases in the mean terminal voltages as large as 31 percent were measured with the applied magnetic fields. These effects are shown to be associated with the increased current conduction along the applied magnetic field lines in the near-anode region. These results also suggest a reduction in frequency and intensity of current-concentrating filaments and anode spots, which contribute to erosion. Overall, both applied

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

    Science.gov (United States)

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

    2017-07-01

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

  6. Performance of a Permanent-Magnet Cylindrical Hall-Effect Thruster

    Science.gov (United States)

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

    2009-01-01

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

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

    Directory of Open Access Journals (Sweden)

    F.A. Ali

    2014-12-01

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

  8. Investigation of Low Discharge Voltage Hall Thruster Operating Modes and Ionization Processes

    Science.gov (United States)

    2009-08-14

    a null-type, inverted pendulum thrust stand based on the NASA GRC design.11 The thruster is shown mounted to the thrust stand in Figure 3... cloud of neutral propellant. This thruster operation was studied in detail using the far-field diagnostics and characterized with variations in

  9. A Robust Digital Autopilot for Spacecraft Equipped with Pulse-Operated Thrusters

    Science.gov (United States)

    Thurman, S. W.; Flashner, H.

    1996-01-01

    The analysis and design of attitude control systems for spacecraft employing pulse-operated (on-off) thrusters is usually accomplished through a combination of modeling approximations and empirical techniques. In this paper a new thruster pulse-modulation scheme for pointing and tracking applications is developed from nonlinear control theory.

  10. Ion behavior in low-power magnetically shielded and unshielded Hall thrusters

    Science.gov (United States)

    Grimaud, L.; Mazouffre, S.

    2017-05-01

    Magnetically shielded Hall thrusters achieve a longer lifespan than traditional Hall thrusters by reducing wall erosion. The lower erosion rate is attributed to a reduction of the high energy ion population impacting the walls. To investigate this phenomenon, the ion velocity distribution functions are measured with laser induced fluorescence at several points of interest in the magnetically shielded ISCT200-MS and the unshielded ISCT200-US Hall thrusters. The center of the discharge channel is probed to highlight the difference in plasma positioning between the shielded and unshielded thrusters. Erosion phenomena are investigated by taking measurements of the ion velocity distribution near the inner and outer wall as well as above the magnetic poles where some erosion is observed. The resulting distribution functions show a displacement of the acceleration region from inside the channel in the unshielded thruster to downstream of the exit plane in the ISCT200-MS. Near the walls, the unshielded thruster displays both a higher relative ion density as well as a significant fraction of the ions with velocities toward the walls compared to the shielded thruster. Higher proportions of high velocity ions are also observed. Those results are in accordance with the reduced erosion observed. Both shielded and unshielded thrusters have large populations of ions impacting the magnetic poles. The mechanism through which those ions are accelerated toward the magnetic poles has so far not been explained.

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

    Science.gov (United States)

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

    2016-10-01

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

  12. Hall Effect Thruster Interactions Data From the Russian Express-A2 and Express-A3 Satellites

    Science.gov (United States)

    Sitnikova, N.; Volkov, D.; Maximov, I.; Petrusevich, V.; Allen, D.

    2003-01-01

    This 12-part report documents the data obtained from various sensor measurements taken aboard the Russian Express-A2 and Express-A3 spacecraft in Geosynchronous Earth Orbit (GEO). These GEO communications satellites, which were designed and built by NPO Prikladnoy Mekhaniki (NPO PM) of Zheleznogorsk, Russia, utilize Hall thruster propulsion systems for north-south and east-west stationkeeping and as of June 2002, were still operating at 80 E. and 11 W., respectively. Express-A2 was launched on March 12, 2000, while Express-A3 was launched on June 24, 2000. The diagnostic equipment from which these data were taken includes electric field strength sensors, ion current and energy sensors, and pressure sensors. The diagnostics and the Hall thruster propulsion systems are described in detail along with lists of tabular data from those diagnostics and propulsion system and other satellite systems. Space Power, Inc., now part of Pratt & Whitney's Chemical Systems Division, under contract NAS3 99151 to the NASA Glenn Research Center, obtained these data over several periods from March 12, 2000, through September 30, 2001. Each of the 12 individual reports describe, in detail, the propulsion systems as well as the diagnostic sensors utilized. Finally, parts 11 and 12 include the requirements to which NPO PM prepared and delivered these data.

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2005-12-01

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

  14. Electron Cross-field Transport in a Low Power Cylindrical Hall Thruster

    Energy Technology Data Exchange (ETDEWEB)

    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.

  15. Background Pressure Effects on Krypton Hall Effect Thruster Internal Acceleration

    Science.gov (United States)

    2013-08-01

    krypton operation of the BHT -600 at the conditions in Table 2 yields a thrust of 22.4 mN corresponding to an anode efficiency of approximately 31...measurement volume is ap- proximately 500 µm diameter by 1 mm length.   Measurement Domain Figure 3 shows a cross-section of the BHT -600 Hall effect...of the BHT -600 Hall effect thruster with measurement volume shown in red. All dimensions are given in mm.     tion of the transition

  16. Design of automatic thruster assisted mooring systems for ships

    Directory of Open Access Journals (Sweden)

    Jan P. Strand

    1998-04-01

    Full Text Available This paper addresses the mathematical modelling and controller design of an automatic thruster assisted position mooring system. Such control systems are applied to anchored floating production offloading and storage vessels and semi-subs. The controller is designed using model based control with a LQG feedback controller in conjunction with a Kalman filter. The controller design is in addition to the environmental loads accounting for the mooring forces acting on the vessel. This is reflected in the model structure and in the inclusion of new functionality.

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

    Science.gov (United States)

    Prak, Moline K.

    2004-01-01

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

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    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.

  20. 2D particle-in-cell simulations of the electron drift instability and associated anomalous electron transport in Hall-effect thrusters

    Science.gov (United States)

    Croes, Vivien; Lafleur, Trevor; Bonaventura, Zdeněk; Bourdon, Anne; Chabert, Pascal

    2017-03-01

    In this work we study the electron drift instability in Hall-effect thrusters (HETs) using a 2D electrostatic particle-in-cell (PIC) simulation. The simulation is configured with a Cartesian coordinate system modeling the radial-azimuthal (r{--}θ ) plane for large radius thrusters. A magnetic field, {{B}}0, is aligned along the Oy axis (r direction), a constant applied electric field, {{E}}0, along the Oz axis (perpendicular to the simulation plane), and the {{E}}0× {{B}}0 direction is along the Ox axis (θ direction). Although electron transport can be well described by electron–neutral collisions for low plasma densities, at high densities (similar to those in typical HETs), a strong instability is observed that enhances the electron cross-field mobility; even in the absence of electron–neutral collisions. The instability generates high frequency (of the order of MHz) and short wavelength (of the order of mm) fluctuations in both the azimuthal electric field and charged particle densities, and propagates in the {{E}}0× {{B}}0 direction with a velocity close to the ion sound speed. The correlation between the electric field and density fluctuations (which leads to an enhanced electron–ion friction force) is investigated and shown to be directly responsible for the increased electron transport. Results are compared with a recent kinetic theory, showing good agreement with the instability properties and electron transport.

  1. Ion angular distribution simulation of the HEMP Thruster

    Science.gov (United States)

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

    2016-10-01

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

  2. Assessment of Pole Erosion in a Magnetically Shielded Hall Thruster

    Science.gov (United States)

    Mikellides, Ioannis G.; Ortega, Alejandro L.

    2014-01-01

    Numerical simulations of a 6-kW laboratory Hall thruster called H6 have been performed to quantify the erosion rate at the inner pole. The assessments have been made in two versions of the thruster, namely the unshielded (H6US) and magnetically shielded (H6MS) configurations. The simulations have been performed with the 2-D axisymmetric code Hall2De which employs a new multi-fluid ion algorithm to capture the presence of low-energy ions in the vicinity of the poles. It is found that the maximum computed erosion rate at the inner pole of the H6MS exceeds the measured rate of back-sputtered deposits by 4.5 times. This explains only part of the surface roughening that was observed after a 150-h wear test, which covered most of the pole area exposed to the plasma. For the majority of the pole surface the computed erosion rates are found to be below the back-sputter rate and comparable to those in the H6US which exhibited little to no sputtering in previous tests. Possible explanations for the discrepancy are discussed.

  3. Engineering Risk Assessment of Space Thruster Challenge Problem

    Science.gov (United States)

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

    2014-01-01

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

  4. Assessment of Pole Erosion in a Magnetically Shielded Hall Thruster

    Science.gov (United States)

    Mikellides, Ioannis G.; Ortega, Alejandro L.

    2014-01-01

    Numerical simulations of a 6-kW laboratory Hall thruster called H6 have been performed to quantify the erosion rate at the inner pole. The assessments have been made in two versions of the thruster, namely the unshielded (H6US) and magnetically shielded (H6MS) configurations. The simulations have been performed with the 2-D axisymmetric code Hall2De which employs a new multi-fluid ion algorithm to capture the presence of low-energy ions in the vicinity of the poles. It is found that the maximum computed erosion rate at the inner pole of the H6MS exceeds the measured rate of back-sputtered deposits by 4.5 times. This explains only part of the surface roughening that was observed after a 150-h wear test, which covered most of the pole area exposed to the plasma. For the majority of the pole surface the computed erosion rates are found to be below the back-sputter rate and comparable to those in the H6US which exhibited little to no sputtering in previous tests. Possible explanations for the discrepancy are discussed.

  5. Space Shuttle vernier thruster long-life chamber development

    Science.gov (United States)

    Krohn, Douglas D.

    1990-01-01

    The Space Shuttle Reaction Control Subsystem (RCS) vernier thruster is a pressure fed engine that utilizes storable propellants to provide precise attitude control for the Orbiter. The current vernier thruster is life limited due to its chamber material. By developing an iridium-lined rhenium chamber for the vernier, substantial gains could be achieved in the operational life of the chamber. The present RCS vernier, its requirements, operating characteristics, and life limitations are described. The current technology status of iridium-lined rhenium is presented along with a description of the operational life capabilities to be gained from implementing this material into the design of a long life vernier chamber. Discussion of the proposed demonstration program to be performed by the NASA Lyndon B. Johnson Space Center to attain additional insight into the application of this technology to the RCS vernier, includes the technical objectives, approach, and program schedule. The plans for further development and integration with the Orbiter and the Shuttle system are also presented.

  6. Characterization of Vacuum Facility Background Gas Through Simulation and Considerations for Electric Propulsion Ground Testing

    Science.gov (United States)

    Yim, John T.; Burt, Jonathan M.

    2015-01-01

    The background gas in a vacuum facility for electric propulsion ground testing is examined in detail through a series of cold flow simulations using a direct simulation Monte Carlo (DSMC) code. The focus here is on the background gas itself, its structure and characteristics, rather than assessing its interaction and impact on thruster operation. The background gas, which is often incorrectly characterized as uniform, is found to have a notable velocity within a test facility. The gas velocity has an impact on the proper measurement of pressure and the calculation of ingestion flux to a thruster. There are also considerations for best practices for tests that involve the introduction of supplemental gas flows to artificially increase the background pressure. All of these effects need to be accounted for to properly characterize the operation of electric propulsion thrusters across different ground test vacuum facilities.

  7. Erosion Measurements in a Diverging Cusped-Field Thruster (Pre Print)

    Science.gov (United States)

    2012-02-01

    3,000 3,000∗ BHT -200 [15] 200 43.5% 1,287-1,519 >1,700 HT-100 [16] 175 25% [4] 300 [17] 1,500 [17]∗∗ SPT-30 [18] 150 26% [19] 600∗ - SPT-20M [20] ...provide longer lifetimes to low-power thrusters with performance capabilities similar to Hall thrusters. Performance similar to the BHT -200, a...only the DCF and BHT -200 thrusters have the reported capability of operating past 1000 h without exposing components of their magnetic circuit. The

  8. Comparison of Numerical and Experimental Time-Resolved Near-Field Hall Thruster Plasma Properties

    Science.gov (United States)

    2014-03-06

    RESOLVED NEAR-FIELD HALL THRUSTER PLASMA PROPERTIES 807 TABLE I BHT -600 HALL THRUSTER AT NOMINAL XENON OPERATING CONDITIONS AND PERFORMANCE [21] 1.2-m...is a 600 W BHT -600 with a 3.2-mm hollow cathode manufactured by the Busek Company (Natick, MA). This thruster has been studied previously using both...electrostatic probes and various opti- cal diagnostics [17]–[20]. The BHT -600 has an acceleration channel outer radius of 32 mm, inner radius of 24 mm

  9. Effects of Anode Temperature on Working Characteristics and Performance of a Low Power Arcjet Thruster

    Institute of Scientific and Technical Information of China (English)

    PAN Wen-Xia; LI Teng; WU Cheng-Kang

    2009-01-01

    An arc-heated thruster of 130-800 W input power is tested in a vacuum chamber at pressures lower than 20 Pa with argon or H_2-N_2 gas mixture as propellant.The time-dependent arc voltage-current curve,outside-surface temperature of the anode nozzle and the produced thrust of the firing arcjet thruster are measured in situ simultaneously,in order to analyze and evaluate the dependence of thruster working characteristics and output properties,such as specific impulse and thrust efficiency,on nozzle temperature.

  10. Satellite Microwave Communication Signal Degradation Due To Hall Thruster Plasma Plumes

    Science.gov (United States)

    Wiley, J. C.; Hallock, G. A.; Spencer, E. A.; Meyer, J. W.; Loane, J. T.

    2001-10-01

    We have developed a geometric optics vector ray-tracing code, BeamServer, for analyzing the effects of Hall thruster plasma plumes on satellite microwave communication signals. The possible effects include main beam attenuation and squinting, side lobe degradation, and induced cross-polarization. We report on a study of Hall current thruster (HCT) mounting positions on a realistic satellite configuration and a study with a highly shaped reflector. Results indicate HCT signal degradation can occur and should be considered in the satellite design process. Initial results of antenna pattern perturbations due to low frequency plume oscillations driven by thruster instabilities are also given.

  11. Modeling of gas ionization and plasma flow in ablative pulsed plasma thrusters

    Science.gov (United States)

    Huang, Tiankun; Wu, Zhiwen; Liu, Xiangyang; Xie, Kan; Wang, Ningfei; Cheng, Yue

    2016-12-01

    A one-dimensional model to study the gas ionization and plasma flow in ablative pulsed plasma thrusters(APPTs) is established in this paper. The discharge process of the APPT used in the LES-6 satellite is simulated to validate the model. The simulation results for the impulse bit and propellant utilization give values of 29.05 μN s and 9.56%, respectively, which are in good agreement with experimental results. To test the new ionization sub-model, the discharge process of a particular APPT, XPPT-1, is simulated, and a numerical result for the propellant utilization of 62.8% is obtained, which also agrees well with experiment. The gas ionization simulation results indicate that an APPT with a lower average propellant ablation rate and higher average electric field intensity between electrodes should have higher propellant utilization. The plasma density distribution between the electrodes of APPTs can also be obtained using the new model, and the numerical results show that the plasma generation and flow are discontinuous, which is in good agreement with past experimental results of high-speed photography. This model provides a new tool with which to study the physical mechanisms of APPTs and a reference for the design of high-performance APPTs.

  12. Effect of the Discharge Voltage on the Performance of the Hall Thruster

    Science.gov (United States)

    Duan, Ping; Cao, An-Ning; Liu, Guang-Rui; Bian, Xing-Yu; Yin, Yan; Chen, Long

    2015-09-01

    A two-dimensional physical model is established according to the discharging process in the Hall thruster discharge channel. By using particle-in-cell method, the influences of the discharge voltage on the distribution of the potential, ion number density, electron temperature and ion radial velocity are investigated in a fixed magnetic field configuration, and the reason of discharge current increasement along with discharge voltage is also analyzed. It is found that, while the discharge voltage increases during 250-650V, the distribution of electric potential has a small expansion towards anode in axis direction, the ion radial velocity at the exit is reduced, and collision frequency between ions and wall surface is also decreased. Meanwhile, the saturation of electron temperature is observed when discharge voltage is in the range of 400-450V, and the electron temperature decreases. When the discharge voltage increases to 700V, the distribution of potential expands towards anode in axis direction significantly, the acceleration region length is greatly increased, the ionization region is compressed to the vicinity of anode, the ion radial velocity near the wall increases and collision frequency between ions and wall surface is enhanced. As the increasement of near wall conductive current and high energy electron number inside the discharge channel, the discharge current increases monotonously with the increase of discharge voltage.

  13. Neutral-depletion-induced asymmetric plasma density profile and momentum transport in a helicon thruster

    Science.gov (United States)

    Takahashi, Kazunori; Takao, Yoshinori; Chiba, Aiki; Ando, Akira

    2016-09-01

    Axial momentum lost to a lateral wall of a helicon source is directly measured by using a pendulum force balance, where only the lateral wall is attached to the balance immersed in 60-cm-diam and 1.4-m-long vacuum tank (pumping speed of 300-400 L/s). When operating the source with highly ionized krypton and xenon, the strong density decay along the axis is observed inside the source tube, which seems to be due to the neutral depletion. Under such a condition, a non-negligible loss of the axial momentum to the lateral wall is detected. The presently detected loss of the axial momentum indicates the presence of the ions which are axially accelerated by the electric field in the plasma core and then lost to the lateral wall. Furthermore, the helicon thruster immersed in 1-m-diam and 2-m-long vacuum tank (pumping speed of 4000-5000 L/s) is operated at high rf power up to 5 kW in argon, to demonstrate the neutral-depletion-induced axially asymmetric density profile. Combination between the Langmuir probe and the optical diagnosis indicates that the neutral density at the axial center of the source is reduced to 20% of the initial neutral density. This work is partially supported by grant-in-aid for scientific research (16H04084 and 26247096) from the Japan Society for the Promotion of Science.

  14. A two-dimensional (azimuthal-axial) particle-in-cell model of a Hall thruster

    Energy Technology Data Exchange (ETDEWEB)

    Coche, P.; Garrigues, L., E-mail: laurent.garrigues@laplace.univ-tlse.fr [LAPLACE (Laboratoire Plasma et Conversion d' Energie), Université de Toulouse, UPS, INPT Toulouse 118, route de Narbonne, F-31062 Toulouse cedex 9 (France); CNRS, LAPLACE, F-31062 Toulouse (France)

    2014-02-15

    We have developed a two-dimensional Particle-In-Cell model in the azimuthal and axial directions of the Hall thruster. A scaling method that consists to work at a lower plasma density to overcome constraints on time-step and grid-spacing is used. Calculations are able to reproduce the breathing mode due to a periodic depletion of neutral atoms without the introduction of a supplementary anomalous mechanism, as in fluid and hybrid models. Results show that during the increase of the discharge current, an electron-cyclotron drift instability (frequency in the range of MHz and wave number on the order of 3000 rad s{sup −1}) is formed in the region of the negative gradient of magnetic field. During the current decrease, an axial electric wave propagates from the channel toward the exhaust (whose frequency is on the order of 400 kHz) leading to a broadening of the ion energy distribution function. A discussion about the influence of the scaling method on the calculation results is also proposed.

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

    Science.gov (United States)

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

    1997-01-01

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

  16. Power Processing for a Conceptual Project Prometheus Electric Propulsion System

    Science.gov (United States)

    Scina, Joseph E., Jr.; Aulisio, Michael; Gerber, Scott S.; Hewitt, Frank; Miller, Leonard; Elbuluk, Malik; Pinero, Luis R. (Technical Monitor)

    2005-01-01

    NASA has proposed a bold mission to orbit and explore the moons of Jupiter. This mission, known as the Jupiter Icy Moons Orbiter (JIMO), would significantly increase NASA s capability to explore deep space by making use of high power electric propulsion. One electric propulsion option under study for JIMO is an ion propulsion system. An early version of an ion propulsion system was successfully used on NASA's Deep Space 1 mission. One concept for an ion thruster system capable of meeting the current JIMO mission requirement would have individual thrusters that are 16 to 25 kW each and require voltages as high as 8.0 kV. The purpose of this work is to develop power processing schemes for delivering the high voltage power to the spacecraft ion thrusters based upon a three-phase AC distribution system. In addition, a proposed DC-DC converter topology is presented for an ion thruster ancillary supply based upon a DC distribution system. All specifications discussed in this paper are for design convenience and are speculative in nature.

  17. Three-dimensional particle simulation of back-sputtered carbon in electric propulsion test facility

    Science.gov (United States)

    Zheng, Hongru; Cai, Guobiao; Liu, Lihui; Shang, Shengfei; He, Bijiao

    2017-03-01

    The back-sputtering deposition on thruster surface caused by ion bombardment on chamber wall material affects the performance of thrusters during the ground based electric propulsion endurance tests. In order to decrease the back-sputtering deposition, most of vacuum chambers applied in electric propulsion experiments are equipped with anti-sputtering targets. In this paper, a three-dimensional model of plume experimental system (PES) including double layer anti-sputtering target is established. Simulation cases are made to simulate the plasma environment and sputtering effects when an ion thruster is working. The particle in cell (PIC) method and direct simulation Monte Carlo (DSMC) method is used to calculate the velocity and position of particles. Yamamura's model is used to simulate the sputtering process. The distribution of sputtered anti-sputtering target material is presented. The results show that the double layer anti-sputtering target can significantly reduce the deposition on thruster surface. The back-sputtering deposition rates on thruster exit surface for different cases are compared. The chevrons on the secondary target are rearranged to improve its performance. The position of secondary target has relation with the ion beam divergence angle, and the radius of the vacuum chamber. The back-sputtering deposition rate is lower when the secondary target covers the entire ion beam.

  18. Electric Propulsion Plume Simulations Using Parallel Computer

    Directory of Open Access Journals (Sweden)

    Joseph Wang

    2007-01-01

    Full Text Available A parallel, three-dimensional electrostatic PIC code is developed for large-scale electric propulsion simulations using parallel supercomputers. This code uses a newly developed immersed-finite-element particle-in-cell (IFE-PIC algorithm designed to handle complex boundary conditions accurately while maintaining the computational speed of the standard PIC code. Domain decomposition is used in both field solve and particle push to divide the computation among processors. Two simulations studies are presented to demonstrate the capability of the code. The first is a full particle simulation of near-thruster plume using real ion to electron mass ratio. The second is a high-resolution simulation of multiple ion thruster plume interactions for a realistic spacecraft using a domain enclosing the entire solar array panel. Performance benchmarks show that the IFE-PIC achieves a high parallel efficiency of ≥ 90%

  19. Guide to Flow Measurement for Electric Propulsion Systems

    Science.gov (United States)

    Frieman, Jason D.; Walker, Mitchell L. R.; Snyder, Steve

    2013-01-01

    In electric propulsion (EP) systems, accurate measurement of the propellant mass flow rate of gas or liquid to the thruster and external cathode is a key input in the calculation of thruster efficiency and specific impulse. Although such measurements are often achieved with commercial mass flow controllers and meters integrated into propellant feed systems, the variability in potential propellant options and flow requirements amongst the spectrum of EP power regimes and devices complicates meter selection, integration, and operation. At the direction of the Committee on Standards for Electric Propulsion Testing, a guide was jointly developed by members of the electric propulsion community to establish a unified document that contains the working principles, methods of implementation and analysis, and calibration techniques and recommendations on the use of mass flow meters in laboratory and spacecraft electric propulsion systems. The guide is applicable to EP devices of all types and power levels ranging from microthrusters to high-power ion engines and Hall effect thrusters. The establishment of a community standard on mass flow metering will help ensure the selection of the proper meter for each application. It will also improve the quality of system performance estimates by providing comprehensive information on the physical phenomena and systematic errors that must be accounted for during the analysis of flow measurement data. This paper will outline the standard methods and recommended practices described in the guide titled "Flow Measurement for Electric Propulsion Systems."

  20. Multi-Scale Modeling of Novel Hall Thrusters: Understanding Physics of CHT and DCF Thrusters

    Science.gov (United States)

    2011-12-30

    benefit of being able to capture not just a general non - Maxwellian velocity distribution space , but also radial variation in mass flux. B. Hybrid...continues to grow as more ions are accelerated from the bulk plasma towards the wall. This result is somewhat non -physical, since in a real device, the...The QN solution is non -physical, since the magnitude of the near-wall electric field is directly related to the cell spacing . The wall effect in the

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

    Science.gov (United States)

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

    2015-01-01

    The design and testing of a small inductive pulsed plasma thruster (IPPT), shown in Fig. 1 with all the major subsystems required for a thruster of this kind are described. Thrust measurements and imaging of the device operated in rep-rated mode are presented to quantify the performance envelope of the device. The small IPPT described in this paper was designed to serve as a test-bed for the pulsed gas-valves and solid-state switches required for a IPPTs. A modular design approach was used to permit future modifications and upgrades. The thruster consists of the following sub-systems: a) a multi-turn, spiral-wound acceleration coil (27 cm o.d., 10 cm i.d.) driven by a 10 microFarad capacitor and switched with a high-voltage thyristor, b) a fast pulsed gas-valve, and c.) a glow-discharge pre-ionizer (PI) circuit. The acceleration-coil circuit may be operated at voltages up to 4 kV (the thyristor limit is 4.5 kV). The device may be operated at rep-rates up to 30 Hz with the present gas-valve. Thrust measurements and imaging of the device operated in rep-rated mode will be presented. The pre-ionizer consists of a 0.3 microFarad capacitor charged to 4 kV and connected to two annular stainless-steel electrodes bounding the area of the coil-face. The 4 kV potential is held across them and when the gas is puffed in over the coil, the PI circuit is completed, and a plasma is formed. Even at the less than optimal base-pressure in the chamber (approximately 5 × 10(exp -4) torr), the PI held-off the applied voltage, and only discharged upon command. For a capacitor charge of 2 kV the peak coil current is 4.1 kA, and during this pulse a very bright discharge (much brighter than from the PI alone) was observed (see Fig. 2). Interestingly, for discharges at this charge voltage the PI was not required as the current rise rate, dI/dt, of the coil itself was sufficient to ionize the gas.

  2. HIGH ENERGY REPLACEMENT FOR TEFLON PROPELLANT IN PULSED PLASMA THRUSTERS Project

    Data.gov (United States)

    National Aeronautics and Space Administration — This program will utilize a well-characterized Pulsed Plasma Thruster (PPT) to test experimental high-energy extinguishable solid propellants (HE), instead of...

  3. Colloid Thruster for Attitude Control Systems (ACS) and Tip-off Control Applications Project

    Data.gov (United States)

    National Aeronautics and Space Administration — We propose to develop and test key technologies needed for an integrated, high thrust colloid thruster system with no moving parts, for spacecraft attitude control...

  4. Feasibility of a 5mN Laser-Driven Mini-Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — We have developed a next-generation thruster under a Phase II SBIR which we believe can meet NASA requirements after some modifications and improvements. It is the...

  5. Kinetic Molecular Dynamic Model of Hall Thruster Channel Wall Erosion Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Hall thrusters are being considered for many space missions because their high specific impulse delivers a larger payload mass fraction than chemical rockets. With a...

  6. Simulation of a Cold Gas Thruster System and Test Data Correlation

    Science.gov (United States)

    Hauser, Daniel M.; Quinn, Frank D.

    2012-01-01

    During developmental testing of the Ascent Abort 1 (AA-1) cold gas thruster system, unexpected behavior was detected. Upon further review the design as it existed may not have met the requirements. To determine the best approach for modifying the design, the system was modeled with a dynamic fluid analysis tool (EASY5). The system model consisted of the nitrogen storage tank, pressure regulator, thruster valve, nozzle, and the associated interconnecting line lengths. The regulator and thruster valves were modeled using a combination of the fluid and mechanical modules available in EASY5. The simulation results were then compared against actual system test data. The simulation results exhibited behaviors similar to the test results, such as the pressure regulators response to thruster firings. Potential design solutions were investigated using the analytical model parameters, including increasing the volume downstream of the regulator and increasing the orifice area. Both were shown to improve the regulator response.

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

    Science.gov (United States)

    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.

  8. Effect of Ambipolar Potential on the Propulsive Performance of the GDM Plasma Thruster Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The gasdynamic mirror (GDM) plasma thruster has the ability to confine high-density plasma for the length of time required to heat it to the temperatures...

  9. Wide Throttling, High Throughput Hall Thruster for Science and Exploration Missions Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In response to Topic S3.04 "Propulsion Systems," Busek Co. Inc. will develop a high throughput Hall effect thruster with a nominal peak power of 1-kW and wide...

  10. Wide Throttling, High Throughput Hall Thruster for Science and Exploration Missions Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In response to Topic S3-04 "Propulsion Systems," Busek proposes to develop a high throughput Hall effect thruster with a nominal peak power of 1-kW and wide...

  11. Low Cost Refractory Matrix Composite Thruster for Mars Ascent Vehicles Project

    Data.gov (United States)

    National Aeronautics and Space Administration — The long-term goal for this effort is to develop a low-cost, high-temperature thruster. Within the attitude control propulsion community, many efforts have focused...

  12. Lifetime Improvement of Large Scale Green Monopropellant Thrusters via Novel, Long-Life Catalysts Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Busek proposes to develop a high performance, non-toxic storable, "green" monopropellant thruster suitable for in-space reaction control propulsion. The engine will...

  13. Propellantless Spacecraft Formation-Flying and Maneuvering with Photonic Laser Thrusters Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Until the former NIAC was closed, we had investigated a nano-meter accuracy formation flight method based on photon thrusters and tethers, Photon Tether Formation...

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

    Science.gov (United States)

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

    2015-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

    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.

  17. Plasma Reactors and Plasma Thrusters Modeling by Ar Complete Global Models

    Directory of Open Access Journals (Sweden)

    Chloe Berenguer

    2012-01-01

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

  18. Experimental Investigation of Two Interacting Thruster-Plumes Downstream of the Nozzles

    OpenAIRE

    Holz, André; Dettleff, Georg; Hannemann, Klaus; Ziegenhagen, Stefan

    2012-01-01

    The plume-plume interaction of two small cold gas thrusters is investigated under high vacuum conditions in the DLR high vacuum plume test facility STG-CT. In this paper we concentrate on the interaction downstream of the nozzles. After introducing the experimental equipment, characteristics of shock interaction are presented. Furthermore the appropriateness of the Penetration Knudsen Number for predicting the type of interaction also for thruster plumes is investigated.

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

    Science.gov (United States)

    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.

  20. Performance of a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets

    Science.gov (United States)

    Polzin, Kurt A.; Raitses, Yevgeny; Fisch, Nathaniel J.

    2008-01-01

    While Hall thrusters can operate at high efficiency at kW power levels, it is difficult to construct one that operates over a broad envelope down to 100W 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. An alternative approach is to employ a cylindrical Hall thruster (CHT) geometry. Laboratory model CHTs have operated at power levels ranging from the order of 50 Watts up to 1 kW. These thrusters exhibit performance characteristics which are comparable to conventional, annular Hall thrusters of similar size. Compared to the annular Hall thruster, the CHT has a lower insulator surface area to discharge chamber volume ratio. 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. This makes the CHT geometry promising for low-power applications. Recently, a CHT that uses permanent magnets to produce the magnetic field topology was tested. This thruster has the promise of reduced power consumption over previous CHT iterations that employed electromagnets. Data are presented for two purposes: to expose the effect different controllable parameters have on the discharge and to summarize performance measurements (thrust, Isp, efficiency) obtained using a thrust stand. These data are used to gain insight into the thruster's operation and to allow for quantitative comparisons between the permanent magnet CHT and the electromagnet CHT.

  1. Hafnia-rich mixed oxide ceramics of the system HfO2-ZrO2-TiO2 for heaters and heat exchangers in electrothermal thrusters: The effects of titania on selected electrical and mechanical properties of Hafnia-rich mixed oxides in the system Hafnia-Zirconia-Titania, volume 1

    Science.gov (United States)

    Staszak, Paul Russell; Wirtz, G. P.; Berg, M.; Brown, S. D.

    1988-01-01

    A study of the effects of titania on selected properties of hafnia-rich mixed oxides in the system hafnia-zirconia-titania (HZT) was made in the region 5 to 20 mol percent titania. The studied properties included electrical conductivity, thermal expansion, and fracture strength and toughness. The effects of titania on the properties were studied for the reduced state as well as the oxidized state of the sintered mixed oxides. X-ray analysis showed that the materials were not always single phase. The oxidized compositions went from being monoclinic solid solutions at low titania additions to having three phases (two monoclinic and a titanate phase) at high additions of titania. The reduced compositions showed an increasing cubic phase presence mixed with the monoclinic phase as titania was added. The electrical conductivity increased with temperature at approximately 0.1 mhos/cm at 1700 C for all compositions. The thermal expansion coefficient decreased with increasing titania as did the monoclinic to tetragonal transformation temperature. The fracture strength of the oxidized bars tended to decrease with the addition of titania owing to the presence of the second phase titania. The fracture strength of the reduced bars exhibited a minimum corresponding to a two-phase region of monoclinic and cubic phases. When the second phases were suppressed, the titania tended to increase the fracture strength slightly in both the oxidized and reduced states. The fracture toughness followed similar trends.

  2. Elimination of Lifetime Limiting Mechanism of Hall Thrusters

    Science.gov (United States)

    Jacobson, David T. (Inventor); Manzella, David H. (Inventor)

    2009-01-01

    A Hall thruster includes inner and outer electromagnets, with the outer electromagnet circumferentially surrounding the inner electromagnet along a centerline axis and separated therefrom, inner and outer poles, in physical connection with their respective inner and outer electromagnets, with the inner pole having a mostly circular shape and the outer pole having a mostly annular shape, a discharge chamber separating the inner and outer poles, a combined anode electrode/gaseous propellant distributor, located at an upstream portion of the discharge chamber and supplying propellant gas and an actuator, in contact with a sleeve portion of the discharge chamber. The actuator is configured to extend the sleeve portion or portions of the discharge chamber along the centerline axis with respect to the inner and outer poles.

  3. Study of Plume Characteristics of a Stationary Plasma Thruster

    Institute of Scientific and Technical Information of China (English)

    QIAN Zhong; WANG Pingyang; DU Zhaohui; KANG Xiaolu

    2008-01-01

    Electron density and temperature of the plume are measured by a double Langmuir probe in an experimental chamber.A numerical model based on both particle-in-cell scheme and direct simulation Monte Carlo hybrid method is developed to simulate the flow field of plume.The equation for plasma potential is solved by alternative direction implicit technique. The simulation is verified by comparing the computational results with the measured data.The study indicates that the electron temperature of flow field is about 2 eV and the electron density is about 2.5 × 1016 ~ 5 × 1015 m-3 at the central line with a distance of 0.3 ~ 1.0 m downstream of the thruster exit.The model can well predict the flow field parameters of the steady plume.The efforts of this paper are referable for further investigation.

  4. Power electronics for a 1-kilowatt arcjet thruster

    Science.gov (United States)

    Gruber, R. P.

    1986-01-01

    After more than two decades, new space mission requirements have revived interest in arcjet systems. The preliminary development and demonstration of new, high efficiency, power electronic concepts for start up and steady state control of dc arcjets is reported. The design comprises a pulse width modulated power converter which is closed loop configured to give fast current control. An inductor, in series with the arcjet, serves the dual role of providing instantaneous current control, as well as a high voltage arc ignition pulse. Benchmark efficiency, transient response, regulation, and ripple data are presented. Tests with arcjets demonstrate that the power electronics breadboard can start thrusters consistently with no apparent damage and transfer reliably to the nondestructive high voltage arc mode in less than a second.

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

    Science.gov (United States)

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

    2016-01-01

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

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

    KAUST Repository

    Matlock, Taylor

    2010-07-25

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

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

    Science.gov (United States)

    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.

  8. The electric rail gun for space propulsion

    Science.gov (United States)

    Bauer, D. P.; Barber, J. P.; Vahlberg, C. J.

    1981-01-01

    An analytic feasibility investigation of an electric propulsion concept for space application is described. In this concept, quasistatic thrust due to inertial reaction to repetitively accelerated pellets by an electric rail gun is used to propel a spacecraft. The study encompasses the major subsystems required in an electric rail gun propulsion system. The mass, performance, and configuration of each subsystem are described. Based on an analytic model of the system mass and performance, the electric rail gun mission performance as a reusable orbital transfer vehicle (OTV) is analyzed and compared to a 30 cm ion thruster system (BIMOD) and a chemical propulsion system (IUS) for payloads with masses of 1150 kg and 2300 kg. For system power levels in the range from 25 kW(e) to 100 kW(e) an electric rail gun OTV is more attractive than a BIMOD system for low Earth orbit to geosynchronous orbit transfer durations in the range from 20 to 120 days.

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

    Science.gov (United States)

    Polzin, Kurt; Korman, Valentin

    2009-01-01

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

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

    Science.gov (United States)

    Walker, Mitchell Louis Ronald

    2005-07-01

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

  11. Influence of Power System Technology on Electric Propulsion Missions

    Science.gov (United States)

    Oleson, Steven R.

    1995-01-01

    Electric propulsion (EP) thruster technology, with efficient lightweight power systems can provide substantial reductions in propulsion system wet mass due to the high specific impulse (Isp) of the thrusters. Historically, the space power systems are too massive for many potential orbital missions. The objective of this paper is to show the impact of current power system technology on EP mission performance and determine what technology advancements are needed to make EP beneficial for earth orbital applications. The approach of the paper is to model the electric propulsion system and orbital mission using a partial parametric method. Various missions are analyzed from orbit maintenance to orbit transfer. Results portray the relationship between mission performance and power technology level. Conclusions show which mission applications currently have acceptable power technology, and which mission applications require power technology improvements.

  12. Astrium Approach For Plume Flow And Impingement Of 10 N Bipropellant Thruster

    Science.gov (United States)

    Theroude, Christophe; Scremin, G.; Wartelski, Matias

    2011-05-01

    Plume impingement on spacecraft surfaces due to chemical propulsion is a major concern during satellite operations. Indeed, thrusters plume induces disturbing forces and torques, contamination as well as thermal fluxes on sensitive surfaces. These effects, that have to be accurately predicted, influence the satellite design: thrusters orientation, MLI design, instruments protections, etc. In order to implement an efficient process of analysis, Astrium uses a two steps approach: first the thruster undisturbed flow field is computed, then the impingement on spacecraft surfaces is evaluated. In this paper, Plumflow, the Astrium Satellites software for undisturbed thrusters’ plume computation, is presented. This software is made of several modules in order to accurately compute the flow field in the different parts of the plume. A first module computes the chemistry in the chamber, then Navier-Stokes equations are solved inside the nozzle where the flow is continuous. After that a DSMC code is used for the transitional regime near the thruster lip and finally an hybrid TPMC/source-flow method computes the free molecular far flow field. The studied case is the Astrium GmbH 10 N bipropellant thruster. Some comparisons are presented between Plumflow and Professor G.A. Bird DSMC software DS2V and with DLR experimental data. These comparisons have shown very satisfactory results. Finally, aiming at computing plume impingement, the plume flow field generated with Plumflow has been interfaced with Professor G.A. Bird 3D DSMC software DS3V. The plume impingement simulation is performed by introducing the undisturbed flow field at a boundary of DS3V computational domain. It allows us to evaluate thermal flux distribution due to Astrium 10 N thruster on a plate adjacent to the thruster and to compare with the Astrium plume impingement software.

  13. Electric power processing, distribution, management and energy storage

    Science.gov (United States)

    Giudici, R. J.

    1980-01-01

    Power distribution subsystems are required for three elements of the SPS program: (1) orbiting satellite, (2) ground rectenna, and (3) Electric Orbiting Transfer Vehicle (EOTV). Power distribution subsystems receive electrical power from the energy conversion subsystem and provide the power busses rotary power transfer devices, switchgear, power processing, energy storage, and power management required to deliver control, high voltage plasma interactions, electric thruster interactions, and spacecraft charging of the SPS and the EOTV are also included as part of the power distribution subsystem design.

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

    Science.gov (United States)

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

    2009-01-01

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

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

    Science.gov (United States)

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

    2014-01-01

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

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

    Directory of Open Access Journals (Sweden)

    Amelia eGreig

    2015-10-01

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

  17. Research of the Fault Diagnosis Method for the Thruster of AUV Based on Information Fusion

    Science.gov (United States)

    Wang, Yu-Jia; Zhang, Ming-Jun; Wu, Juan

    Aiming at the problem of thruster fault diagnosis of AUV, the motion condition model of AUV based on the improved dynamic recursive Elman neural network, and the performance model of thruster based on the Radial Basis Function network were established. And the fault fusion diagnosis method was proposed according to the overall and local fault detection. Through comparing the output value of motion condition model with the measured value of actual speed and angle, it obtained the overall fault information. Also, it obtained the direct fault information through analyzing the residual which was produced by comparing the output of the performance model with the measured value of the actual voltage and current of the each thruster. According to the decision level information fusion of two kinds of information, it realized the fault diagnosis of thrusters and analyzed the fault degree and reliability. The results of the fault-simulation experiment show that the proposed fault fusion diagnosis method for the thruster of AUV is feasible and effective.

  18. Development of a Micro-Thruster Test Facility which fulfils the LISA requirements

    Science.gov (United States)

    Hey, Franz Georg; Keller, A.; Johann, U.; Braxmaier, C.; Tajmar, M.; Fitzsimons, E.; Weise, D.

    2015-05-01

    In the context of investigations for a sufficient attitude control thruster for LISA, we have developed a thruster test facility which consists of a highly precise thrust balance coupled with plasma diagnostics. In parallel to the test facility development, investigations to downscale a High Efficiency Multistage Plasma Thruster (HEMP-T) are also being carried out. The thruster has been used to demonstrate the measurement capabilities of the facility. The setup allows a parallel operation of all instruments and can also be used for other types of μN propulsion systems including cold gas thrusters. The thrust balance consists of two pendulums. As read out a heterodyne laser interferometer is used. Differential wave front sensing (DWS) enables the measurement of the pendulum tilt which, via suitable calibration using an electrostatic comb, can be converted to a thrust. The whole setup is a symmetric configuration enabling a common-mode rejection of the dominant noise sources (e.g. seismic noise etc.). The thrust balance has a demonstrated precision of 0.1 μN. Based on our unique design, this precision can be attained down to 10-3 Hz. Thus, the measurement setup is especially suitable for characterising the thrust noise of potential eLISA propulsion candidates. We give an overview of the design, the present performance and the future plans.

  19. 5.8kV SiC PiN Diode for Switching of High-Efficiency Inductive Pulsed Plasma Thruster Circuits

    Science.gov (United States)

    Toftul, Alexandra; Polzin, Kurt A.; Hudgins, Jerry L.

    2014-01-01

    Inductive Pulsed Plasma Thruster (IPPT) pulse circuits, such as those needed to operate the Pulsed Inductive Thruster (PIT), are required to quickly switch capacitor banks operating at a period of µs while conducting current at levels on the order of at least 10 kA. [1,2] For all iterations of the PIT to date, spark gaps have been used to discharge the capacitor bank through an inductive coil. Recent availability of fast, high-power solid state switching devices makes it possible to consider the use of semiconductor switches in modern IPPTs. In addition, novel pre-ionization schemes have led to a reduction in discharge energy per pulse for electric thrusters of this type, relaxing the switching requirements for these thrusters. [3,4] Solid state switches offer the advantage of greater controllability and reliability, as well as decreased drive circuit dimensions and mass relative to spark gap switches. The use of solid state devices such as Integrated Gate Bipolar Transistors (IGBTs), Gate Turn-off Thyristors (GTOs) and Silicon-Controlled Rectifiers (SCRs) often involves the use of power diodes. These semiconductor devices may be connected antiparallel to the switch for protection from reverse current, or used to reduce power loss in a circuit by clamping off current ringing. In each case, higher circuit efficiency may be achieved by using a diode that is able to transition, or 'switch,' from the forward conducting state ('on' state) to the reverse blocking state ('off' state) in the shortest amount of time, thereby minimizing current ringing and switching losses. Silicon Carbide (SiC) PiN diodes offer significant advantages to conventional fast-switching Silicon (Si) diodes for high power and fast switching applications. A wider band gap results in a breakdown voltage 10 times that of Si, so that a SiC device may have a thinner drift region for a given blocking voltage. [5] This leads to smaller, lighter devices for high voltage applications, as well as reduced

  20. The ISR Asymmetrical Capacitor Thruster: Experimental Results and Improved Designs

    Science.gov (United States)

    Canning, Francis X.; Cole, John; Campbell, Jonathan; Winet, Edwin

    2004-01-01

    A variety of Asymmetrical Capacitor Thrusters has been built and tested at the Institute for Scientific Research (ISR). The thrust produced for various voltages has been measured, along with the current flowing, both between the plates and to ground through the air (or other gas). VHF radiation due to Trichel pulses has been measured and correlated over short time scales to the current flowing through the capacitor. A series of designs were tested, which were increasingly efficient. Sharp features on the leading capacitor surface (e.g., a disk) were found to increase the thrust. Surprisingly, combining that with sharp wires on the trailing edge of the device produced the largest thrust. Tests were performed for both polarizations of the applied voltage, and for grounding one or the other capacitor plate. In general (but not always) it was found that the direction of the thrust depended on the asymmetry of the capacitor rather than on the polarization of the voltage. While no force was measured in a vacuum, some suggested design changes are given for operation in reduced pressures.

  1. Vacuum Plasma Spray (VPS) Material Applications for Thruster Components

    Science.gov (United States)

    Elam, Sandra; Holmes, Richard; Hickman, Robert

    2006-01-01

    A variety of vacuum plasma spray (VPS) material systems have been successfully applied to injector and thrust chamber components. VPS offers a versatile fabrication process with relatively low costs to produce near net shape parts. The materials available with VPS increase operating margins and improve component life by providing superior thermal and oxidation protection in specific engine environments. Functional gradient materials (FGM) formed with VPS allow thrust chamber liners to be fabricated with GRCop-84 (an alloy of copper, chrome, and niobium) and a protective layer of NiCrAlY on the hot wall. A variety of thrust chamber liner designs have been fabricated to demonstrate the versatility of the process. Hot-fire test results have confined the improved durability and high temperature performance of the material systems for thrust chamber liners. Similar FGM s have been applied to provide superior thermal protection on injector faceplates with NiCrAlY and zirconia coatings. The durability of the applied materials has been demonstrated with hot-fire cycle testing on injector faceplates in high temperature environments. The material systems can benefit the components used in booster and main engine propulsion systems. More recent VPS efforts are focused on producing rhenium based material systems for high temperature applications to benefit in-space engines like reaction control system (RCS) thrusters.

  2. Thrust Stand Characterization of the NASA Evolutionary Xenon Thruster (NEXT)

    Science.gov (United States)

    Diamant, Kevin D.; Pollard, James E.; Crofton, Mark W.; Patterson, Michael J.; Soulas, George C.

    2010-01-01

    Direct thrust measurements have been made on the NASA Evolutionary Xenon Thruster (NEXT) ion engine using a standard pendulum style thrust stand constructed specifically for this application. Values have been obtained for the full 40-level throttle table, as well as for a few off-nominal operating conditions. Measurements differ from the nominal NASA throttle table 10 (TT10) values by 3.1 percent at most, while at 30 throttle levels (TLs) the difference is less than 2.0 percent. When measurements are compared to TT10 values that have been corrected using ion beam current density and charge state data obtained at The Aerospace Corporation, they differ by 1.2 percent at most, and by 1.0 percent or less at 37 TLs. Thrust correction factors calculated from direct thrust measurements and from The Aerospace Corporation s plume data agree to within measurement error for all but one TL. Thrust due to cold flow and "discharge only" operation has been measured, and analytical expressions are presented which accurately predict thrust based on thermal thrust generation mechanisms.

  3. Oxygen-hydrogen thrusters for Space Station auxiliary propulsion systems

    Science.gov (United States)

    Berkman, D. K.

    1984-01-01

    The feasibility and technology requirements of a low-thrust, high-performance, long-life, gaseous oxygen (GO2)/gaseous hydrogen (GH2) thruster were examined. Candidate engine concepts for auxiliary propulsion systems for space station applications were identified. The low-thrust engine (5 to 100 lb sub f) requires significant departure from current applications of oxygen/hydrogen propulsion technology. Selection of the thrust chamber material and cooling method needed or long life poses a major challenge. The use of a chamber material requiring a minimum amount of cooling or the incorporation of regenerative cooling were the only choices available with the potential of achieving very high performance. The design selection for the injector/igniter, the design and fabrication of a regeneratively cooled copper chamber, and the design of a high-temperature rhenium chamber were documented and the performance and heat transfer results obtained from the test program conducted at JPL using the above engine components presented. Approximately 115 engine firings were conducted in the JPL vacuum test facility, using 100:1 expansion ratio nozzles. Engine mixture ratio and fuel-film cooling percentages were parametrically investigated for each test configuration.

  4. High Power Nuclear Electric Propulsion (NEP) for Cargo and Propellant Transfer Missions in Cislunar Space

    Science.gov (United States)

    Falck, Robert D.; Borowski, Stanley K.

    2003-01-01

    The performance of Nuclear Electric Propulsion (NEP) in transporting cargo and propellant from Low Earth Orbit (LEO) to the first Earth-Moon Lagrange point (EML1) is examined. The baseline NEP vehicle utilizes a fission reactor system with Brayton power conversion for electric power generation to power multiple liquid hydrogen magnetoplasmadynamic (MPD) thrusters. Vehicle characteristics and performance levels are based on technology availability in a fifteen to twenty year timeframe. Results of numerical trajectory analyses are also provided.

  5. Estimate of Lifetime of Ion Thruster Optics Based on Particle Simulation

    Institute of Scientific and Technical Information of China (English)

    LIU Chang; TANG Haibin; ZHANG Zhenpeng; GU Zuo; LIU Yu

    2008-01-01

    A three-dimensional particle simulation of ion thruster optics with charge-exchange collision was developed in this study. The simulation code was based on tracking ions using the particle-in-cell method, and the Monte Carlo technique was used to model the charge-exchange collision. Simulations were performed for a 20 cm ion thruster optics. The results were compared with the corresponding experimental data from a test of the ion thruster optics for a duration of 800 hours. The Depth-From-Focus (DFF) method was used to measure the erosion depth of the downstream surface of the accelerator grid. The predicted erosion depth of the accelerator grid was consistent reasonably with the corresponding experimental data. The simulation results showed that the accelerator grid would be burned through after 1333 hours.

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

    Science.gov (United States)

    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.

  7. Prediction and experimental measurement of the electromagnetic thrust generated by a microwave thruster system

    Institute of Scientific and Technical Information of China (English)

    Yang Juan; Wang Yu-Quan; Ma Yan-Jie; Li Peng-Fei; Yang Le; Wang Yang; He Guo-Qiang

    2013-01-01

    A microwave thruster system that can convert microwave power directly to thrust without a gas propellant is developed.In the system,a cylindrical tapered resonance cavity and a magnetron microwave source are used respectively as the thruster cavity and the energy source to generate the electromagnetic wave.The wave is radiated into and then reflected from the cavity to form a pure standing wave with non-uniform electromagnetic pressure distribution.Consequently,a net electromagnetic thrust exerted on the axis of the thruster cavity appears,which is demonstrated through theoretical calculation based on the electromagnetic theory.The net electromagnetic thrust is also experimentally measured in the range from 70 mN to 720 mN when the microwave output power is from 80 W to 2500 W.

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

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

    2012-01-01

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

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

    Science.gov (United States)

    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 high-speed camera and a set of high-speed Langmuir probes were implemented to study the effect of varying facility background pressure on thruster operation. 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 accelerationionization 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.

  11. Experimental study of a low-thrust measurement system for thruster ground tests.

    Science.gov (United States)

    Gong, Jingsong; Hou, Lingyun; Zhao, Wenhua

    2014-03-01

    The development of thrusters used for the control of position and orbit of micro-satellites requires thrust stands that can measure low thrust. A new method to measure low thrust is presented, and the measuring device is described. The test results show that the thrust range is up to 1000 mN, the measurement error of the device is lower than ±1% of full scale, and the drift of the zero offset is less than ±1% of full scale. Its response rise time is less than 15 ms. It is employed to measure the working process of a model chemical thruster with repeatability.

  12. Standard Practices for Usage of Inductive Magnetic Field Probes with Application to Electric Propulsion Testing

    Science.gov (United States)

    Polzin, Kurt A.; Hill, Carrie S.

    2013-01-01

    Inductive magnetic field probes (also known as B-dot probes and sometimes as B-probes or magnetic probes) are useful for performing measurements in electric space thrusters and various plasma accelerator applications where a time-varying magnetic field is present. Magnetic field probes have proven to be a mainstay in diagnosing plasma thrusters where changes occur rapidly with respect to time, providing the means to measure the magnetic fields produced by time-varying currents and even an indirect measure of the plasma current density through the application of Ampère's law. Examples of applications where this measurement technique has been employed include pulsed plasma thrusters and quasi-steady magnetoplasmadynamic thrusters. The Electric Propulsion Technical Committee (EPTC) of the American Institute of Aeronautics and Astronautics (AIAA) was asked to assemble a Committee on Standards (CoS) for Electric Propulsion Testing. The assembled CoS was tasked with developing Standards and Recommended Practices for various diagnostic techniques used in the evaluation of plasma thrusters. These include measurements that can yield either global information related to a thruster and its performance or detailed, local data related to the specific physical processes occurring in the plasma. This paper presents a summary of the standard, describing the preferred methods for fabrication, calibration, and usage of inductive magnetic field probes for use in diagnosing plasma thrusters. Inductive magnetic field probes (also called B-dot probes throughout this document) are commonly used in electric propulsion (EP) research and testing to measure unsteady magnetic fields produced by time-varying currents. The B-dot probe is relatively simple in construction, and requires minimal cost, making it a low-cost technique that is readily accessible to most researchers. While relatively simple, the design of a B-dot probe is not trivial and there are many opportunities for errors in

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

    Science.gov (United States)

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

    2010-01-01

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

  14. Preliminary Results of Performance Measurements on a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets

    Science.gov (United States)

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

    2008-01-01

    The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic configurations. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying a higher thrust efficiency. Preliminary thruster performance measurements on this configuration were obtained over a power range of 100-250 W. The thrust levels over this power range were 3.5-6.5 mN, with anode efficiencies and specific impulses spanning 14-19% and 875- 1425 s, respectively. The magnetic field in the thruster was lower for the thrust measurements than the plasma probe measurements due to heating and weakening of the permanent magnets, reducing the maximum field strength from 2 kG to roughly 750-800 G. The discharge current levels observed during thrust stand testing were anomalously high compared to those levels measured in previous experiments with this thruster.

  15. Radioisotope Electric Propulsion (REP): A Near-Term Approach to Nuclear Propulsion

    Science.gov (United States)

    Schmidt, George R.; Manzella, David H.; Kamhawi, Hani; Kremic, Tibor; Oleson, Steven R.; Dankanich, John W.; Dudzinski, Leonard A.

    2009-01-01

    Studies over the last decade have shown radioisotope-based nuclear electric propulsion to be enhancing and, in some cases, enabling for many potential robotic science missions. Also known as radioisotope electric propulsion (REP), the technology offers the performance advantages of traditional reactor-powered electric propulsion (i.e., high specific impulse propulsion at large distances from the Sun), but with much smaller, affordable spacecraft. Future use of REP requires development of radioisotope power sources with system specific powers well above that of current systems. The US Department of Energy and NASA have developed an advanced Stirling radioisotope generator (ASRG) engineering unit, which was subjected to rigorous flight qualification-level tests in 2008, and began extended lifetime testing later that year. This advancement, along with recent work on small ion thrusters and life extension technology for Hall thrusters, could enable missions using REP sometime during the next decade.

  16. Solar-electric-propulsion cargo vehicles for split/sprint Mars mission

    Science.gov (United States)

    Callaghan, Christopher E.; Crowe, Michael D.; Swis, Matthew J.; Mickney, Marcus R.; Montgomery, C. Keith; Walters, Robert; Thoden, Scott

    1991-01-01

    In support of the proposed exploration of Mars, an unmanned cargo ferry SEMM1 (Solar Electric Mars Mission) was designed. The vehicle is based on solar electric propulsion, and required to transport a cargo of 61,000 kg. The trajectory is a combination of spirals; first, out from LEO, then around the sun, then spiral down to low Mars orbit. The spacecraft produces 3.03 MWe power using photovoltaic flexible blanket arrays. Ion thrusters using argon as a propellant were selected to drive the ship, providing about 60 Newtons of thrust in low Earth orbit. The configuration is based on two long truss beams to which the 24 individual, self-deployable, solar arrays are attached. The main body module supports the two beams and houses the computers, electrical, and control equipment. The thruster module is attached to the rear of the main body, and the cargo to the front.

  17. Theory and Numerical Simulation of Plasma-wall Interactions in Electric Propulsion

    Science.gov (United States)

    Mikellides, Ioannis

    2016-10-01

    Electric propulsion (EP) can be an enabling technology for many science missions considered by NASA because it can produce high exhaust velocities, which allow for less propellant mass compared to typical chemical systems. Over the last decade two EP technologies have emerged as primary candidates for several proposed science missions, mainly due to their superior performance and proven record in space flight: the Ion and Hall thrusters. As NASA looks ahead to increasingly ambitious science goals, missions demand higher endurance from the propulsion system. So, by contrast to the early years of development of these thrusters, when the focus was on performance, considerable focus today is shifting towards extending their service life. Considering all potentially life-limiting mechanisms in Ion and Hall thrusters two are of primary concern: (a) the erosion of the acceleration channel in Hall thrusters and (b) the erosion of the hollow cathode. The plasma physics leading to material wear in these devices are uniquely challenging. For example, soon after the propellant is introduced into the hollow cathode it becomes partially ionized as it traverses a region of electron emission. Electron emission involves highly non-linear boundary conditions. Also, the sheath size is typically many times smaller than the characteristic physical scale of the device, yet energy gained by ions through the sheath must be accounted for in the erosion calculations. The plasma-material interactions in Hall thruster channels pose similar challenges that are further exacerbated by the presence of a strong applied magnetic field. In this presentation several complexities associated with plasma-wall interactions in EP will be discussed and numerical simulation results of key plasma properties in two examples, Hall thrusters and hollow cathodes, will be presented.

  18. Space Technology: Game Changing Development Deep Space Engine (DSE) 100 lbf and 5 lbf Thruster Development and Qualification

    Science.gov (United States)

    Barnett, Gregory

    2017-01-01

    Science mission studies require spacecraft propulsion systems that are high-performance, lightweight, and compact. Highly matured technology and low-cost, short development time of the propulsion system are also very desirable. The Deep Space Engine (DSE) 100-lbf thruster is being developed to meet these needs. The overall goal of this game changing technology project is to qualify the DSE thrusters along with 5-lbf attitude control thrusters for space flight and for inclusion in science and exploration missions. The aim is to perform qualification tests representative of mission duty cycles. Most exploration missions are constrained by mass, power and cost. As major propulsion components, thrusters are identified as high-risk, long-lead development items. NASA spacecraft primarily rely on 1960s' heritage in-space thruster designs and opportunities exist for reducing size, weight, power, and cost through the utilization of modern materials and advanced manufacturing techniques. Advancements in MON-25/MMH hypergolic bipropellant thrusters represent a promising avenue for addressing these deficiencies with tremendous mission enhancing benefits. DSE is much lighter and costs less than currently available thrusters in comparable thrust classes. Because MON-25 propellants operate at lower temperatures, less power is needed for propellant conditioning for in-space propulsion applications, especially long duration and/or deep-space missions. Reduced power results in reduced mass for batteries and solar panels. DSE is capable of operating at a wide propellant temperature range (between -22 F and 122 F) while a similar existing thruster operates between 45 F and 70 F. Such a capability offers robust propulsion operation as well as flexibility in design. NASA's Marshall Space Flight Center evaluated available operational Missile Defense Agency heritage thrusters suitable for the science and lunar lander propulsion systems.

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

    Science.gov (United States)

    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

  20. A High-power Electric Propulsion Test Platform in Space

    Science.gov (United States)

    Petro, Andrew J.; Reed, Brian; Chavers, D. Greg; Sarmiento, Charles; Cenci, Susanna; Lemmons, Neil

    2005-01-01

    This paper will describe the results of the preliminary phase of a NASA design study for a facility to test high-power electric propulsion systems in space. The results of this design study are intended to provide a firm foundation for subsequent detailed design and development activities leading to the deployment of a valuable space facility. The NASA Exploration Systems Mission Directorate is sponsoring this design project. A team from the NASA Johnson Space Center, Glenn Research Center, the Marshall Space Flight Center and the International Space Station Program Office is conducting the project. The test facility is intended for a broad range of users including government, industry and universities. International participation is encouraged. The objectives for human and robotic exploration of space can be accomplished affordably, safely and effectively with high-power electric propulsion systems. But, as thruster power levels rise to the hundreds of kilowatts and up to megawatts, their testing will pose stringent and expensive demands on existing Earth-based vacuum facilities. These considerations and the human access to near-Earth space provided by the International Space Station (ISS) have led to a renewed interest in space testing. The ISS could provide an excellent platform for a space-based test facility with the continuous vacuum conditions of the natural space environment and no chamber walls to modify the open boundary conditions of the propulsion system exhaust. The test platform could take advantage of the continuous vacuum conditions of the natural space environment. Space testing would provide open boundary conditions without walls, micro-gravity and a realistic thermal environment. Testing on the ISS would allow for direct observation of the test unit, exhaust plume and space-plasma interactions. When necessary, intervention by on-board personnel and post-test inspection would be possible. The ISS can provide electrical power, a location for

  1. Plume Characterization of a Laboratory Model 22 N GPIM Thruster via High-Frequency Raman Spectroscopy

    Science.gov (United States)

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

    2015-01-01

    The Green Propellant Infusion Mission (GPIM) will demonstrate the capability of a green propulsion system, specifically, one using the monopropellant, AF-M315E. One of the risks identified for GPIM is potential contamination of sensitive areas of the spacecraft from the effluents in the plumes of AF-M315E thrusters. Plume characterization of a laboratory-model 22 N thruster via optical diagnostics was conducted at NASA GRC in a space-simulated environment. A high-frequency pulsed laser was coupled with an electron-multiplied ICCD camera to perform Raman spectroscopy in the near-field, low-pressure plume. The Raman data yielded plume constituents and temperatures over a range of thruster chamber pressures and as a function of thruster (catalyst) operating time. Schlieren images of the near-field plume enabled calculation of plume velocities and revealed general plume structure of the otherwise invisible plume. The measured velocities are compared to those predicted by a two-dimensional, kinetic model. Trends in data and numerical results are presented from catalyst mid-life to end-of-life. The results of this investigation were coupled with the Raman and Schlieren data to provide an anchor for plume impingement analysis presented in a companion paper. The results of both analyses will be used to improve understanding of the nature of AF-M315E plumes and their impacts to GPIM and other future missions.

  2. Power Dependence of the Electron Mobility Profile in a Hall Thruster

    Science.gov (United States)

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

    2014-01-01

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

  3. Propellant Grade Hydrazine in Mono/Bi-propellant Thrusters: Preparation and Performance Evaluation

    Directory of Open Access Journals (Sweden)

    S. Krishnamachary

    2015-03-01

    Full Text Available Propellant grade hydrazine was prepared with 64 per cent yield and 95.5 per cent purity. Purity of the propellant grade hydrazine was determined using wet chemical, gas chromatographic (GC and eudiometric methods. It was observed that the compositions containing blends of hydrazine-methyl alcohol-ammonium nitrate and hydrazine-methyl alcohol-ammonium perchlorate were not found to be frozen even after cooling to -65 °C for 30 minutes. Mono and bi-propellant thrusters were designed and developed to demonstrate the performance of prepared propellant grade hydrazine as a promising rocket fuel. Five static tests with 22 N thruster and one static test with 1 N thruster were performed successfully in mono-propellant mode. The hurdles of chamber pressure oscillations were overcome by compact packing of the catalyst. The desired decomposition and chamber pressure were achieved. One static test was performed successfully with 60 N bi-propellant thruster. The desired chamber pressure and thrust were achieved. The combustion was smooth and C* achieved was higher than that of UH-25, N2O4 combination. The performance of prepared propellant grade hydrazine shows it as a promising rocket fuels.Defence Science Journal, Vol. 65, No. 1, January 2015, pp.31-38, DOI:http://dx.doi.org/10.14429/dsj.65.7986

  4. Revolutionizing Space Propulsion Through the Characterization of Iodine as Fuel for Hall-Effect Thrusters

    Science.gov (United States)

    2011-03-01

    propellant was successfully operated through a BHT -200 thruster in the T6 vacuum facility at Busek Co. Inc. A feed system for the iodine was developed...15  Figure 4: Overall experimental setup of the vacuum ........................................................ 26  Figure 5: BHT -200...and BHC-1500, side view ................................................................. 27  Figure 6: Busek BHT -200 and BHC-1500, front view

  5. A Strategy to Characterize the LISA-Pathfinder Cold Gas Thruster System

    Science.gov (United States)

    Armano, M.; Audley, H.; Auger, G.; Baird, J.; Binetruy, P.; Born, M.; Bortoluzzi, D.; Brandt, N.; Bursi, A.; Caleno, M.; Cavalleri, A.; Cesarini, A.; Cruise, M.; Danzmann, K.; Diepholz, I.; Dolesi, R.; Dunbar, N.; Ferraioli, L.; Ferroni, V.; Fitzsimons, E.; Freschi, M.; Gallegos, J.; Garcia Marirrodriga, C.; Gerndt, R.; Gesa, L. I.; Gibert, F.; Giardini, D.; Giusteri, R.; Grimani, C.; Harrison, I.; Heinzel, G.; Hewitson, M.; Hollington, D.; Hueller, M.; Huesler, J.; Inchauspé, H.; Jennrich, O.; Jetzer, P.; Johlander, B.; Karnesis, N.; Kaune, B.; Korsakova, N.; Killow, C.; Lloro, I.; Maarschalkerweerd, R.; Madden, S.; Mance, D.; Martin, V.; Martin-Porqueras, F.; Mateos, I.; McNamara, P.; Mendes, J.; Mendes, L.; Moroni, A.; Nofrarias, M.; Paczkowski, S.; Perreur-Lloyd, M.; Petiteau, A.; Pivato, P.; Plagnol, E.; Prat, P.; Ragnit, U.; Ramos-Castro, J.; Reiche, J.; Romera Perez, J. A.; Robertson, D.; Rozemeijer, H.; Russano, G.; Sarra, P.; Schleicher, A.; Slutsky, J.; Sopuerta, C. F.; Sumner, T.; Texier, D.; Thorpe, J.; Trenkel, C.; Tu, H. B.; Vitale, S.; Wanner, G.; Ward, H.; Waschke, S.; Wass, P.; Wealthy, D.; Wen, S.; Weber, W.; Wittchen, A.; Zanoni, C.; Ziegler, T.; Zweifel, P.

    2015-05-01

    The cold gas micro-propulsion system that will be used during the LISA-Pathfinder mission will be one of the most important component used to ensure the "free-fall" of the enclosed test masses. In this paper we present a possible strategy to characterize the effective direction and amplitude gain of each of the 6 thrusters of this system.

  6. TRMM Re-Entry Planning: Attitude Determination and Control During Thruster Modes

    Science.gov (United States)

    DeWeese, Keith

    2005-01-01

    The Tropical Rainfall Measuring Mission (TRMM) spacecraft has been undergoing design for a controlled re-entry to Earth. During simulation of the re-entry plan, there was evidence of errors in the attitude determination algorithms during thruster modes. These errors affected the bum efficiency, and thus planning, during re-entry. During thruster modes, the spacecraft attitude is controlled off of integrated Gyro Error Angles that were designed to closely follow the nominal spacecraft pointing frame (Tip Frame). These angles, however, were not exactly mapped to the Tip Frame from the Body Frame. Additionally, in the initial formulation of the thruster mode attitude determination algorithms, several assumptions and approximations were made to conserve processor speed. These errors became noticeable and significant when simulating bums of much longer duration (-10 times) than had been produced in flight. A solution is proposed that uses attitude determination information from a propagated extended Kalman filter that already exists in the TRMM thruster modes. This attitude information is then used to rotate the Gyro Error Angles into the Tip Frame. An error analysis is presented that compares the two formulations. The new algorithm is tested using the TRMM High-Fidelity Simulator and verified with the TRMM Software Testing and Training Facility. Simulation results for both configurations are also presented.

  7. Time-Synchronized Continuous Wave Laser Induced Fluorescence Velocity Measurements of a 600 Watt Hall Thruster

    Science.gov (United States)

    2015-07-01

    used in previous work,21 but well within the linear regime of operation. 3 Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT, Hyogo-Kobe, Japan July...PA Clearance No. 15329 4 BHT-600 Specifications • 600 W annular Hall thruster • Manufactured by Busek Co. • Tested in Chamber 6 at AFRL Nominal

  8. Numerical investigation of two interacting parallel thruster-plumes and comparison to experiment

    Science.gov (United States)

    Grabe, Martin; Holz, André; Ziegenhagen, Stefan; Hannemann, Klaus

    2014-12-01

    Clusters of orbital thrusters are an attractive option to achieve graduated thrust levels and increased redundancy with available hardware, but the heavily under-expanded plumes of chemical attitude control thrusters placed in close proximity will interact, leading to a local amplification of downstream fluxes and of back-flow onto the spacecraft. The interaction of two similar, parallel, axi-symmetric cold-gas model thrusters has recently been studied in the DLR High-Vacuum Plume Test Facility STG under space-like vacuum conditions, employing a Patterson-type impact pressure probe with slot orifice. We reproduce a selection of these experiments numerically, and emphasise that a comparison of numerical results to the measured data is not straight-forward. The signal of the probe used in the experiments must be interpreted according to the degree of rarefaction and local flow Mach number, and both vary dramatically thoughout the flow-field. We present a procedure to reconstruct the probe signal by post-processing the numerically obtained flow-field data and show that agreement to the experimental results is then improved. Features of the investigated cold-gas thruster plume interaction are discussed on the basis of the numerical results.

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

    Energy Technology Data Exchange (ETDEWEB)

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

    2003-09-08

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

  10. Propulsion and control propellers with thruster nozzles primarily for aircraft applications

    Science.gov (United States)

    Pabst, W.

    1986-01-01

    A propulsion and control propeller with thruster nozzles, primarily for aircraft application is described. Adjustability of rotor blades at the hub and pressurized gas expulsion combined with an air propeller increase power. Both characteristics are combined in one simple device, and, furthermore, incorporate overall aircraft control so that mechanisms which govern lateral and horizontal movement become superfluous.

  11. Diagnosis and Fault-Tolerant Control for Thruster-Assisted Position Mooring System

    DEFF Research Database (Denmark)

    Nguyen, Trong Dong; Blanke, Mogens; Sørensen, Asgeir

    2007-01-01

    Development of fault-tolerant control systems is crucial to maintain safe operation of o®shore installations. The objective of this paper is to develop a fault- tolerant control for thruster-assisted position mooring (PM) system with faults occurring in the mooring lines. Faults in line...

  12. Hybrid-PIC Modeling of the Transport of Atomic Boron in a Hall Thruster

    Science.gov (United States)

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

    2015-01-01

    Computational analysis of the transport of boron eroded from the walls of a Hall thruster is performed by implementing sputter yields of hexagonal boron nitride and velocity distribution functions of boron within the hybrid-PIC model HPHall. The model is applied to simulate NASA's HiVHAc Hall thruster at a discharge voltage of 500V and discharge powers of 1-3 kW. The number densities of ground- and 4P-state boron are computed. The density of ground-state boron is shown to be a factor of about 30 less than the plasma density. The density of the excited state is shown to be about three orders of magnitude less than that of the ground state, indicating that electron impact excitation does not significantly affect the density of ground-state boron in the discharge channel or near-field plume of a Hall thruster. Comparing the rates of excitation and ionization suggests that ionization has a greater influence on the density of ground-state boron, but is still negligible. The ground-state boron density is then integrated and compared to cavity ring-down spectroscopy (CRDS) measurements for each operating point. The simulation results show good agreement with the measurements for all operating points and provide evidence in support of CRDS as a tool for measuring Hall thruster erosion in situ.

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

    Directory of Open Access Journals (Sweden)

    Peng Hu

    2016-09-01

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

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

    Science.gov (United States)

    Ceruti, Luca; Polli, Aldo; Galantini, Paolo

    2014-08-01

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

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

    Science.gov (United States)

    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

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

    Science.gov (United States)

    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.

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

    Science.gov (United States)

    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.

  18. ISS Contingency Attitude Control Recovery Method for Loss of Automatic Thruster Control

    Science.gov (United States)

    Bedrossian, Nazareth; Bhatt, Sagar; Alaniz, Abran; McCants, Edward; Nguyen, Louis; Chamitoff, Greg

    2008-01-01

    In this paper, the attitude control issues associated with International Space Station (ISS) loss of automatic thruster control capability are discussed and methods for attitude control recovery are presented. This scenario was experienced recently during Shuttle mission STS-117 and ISS Stage 13A in June 2007 when the Russian GN&C computers, which command the ISS thrusters, failed. Without automatic propulsive attitude control, the ISS would not be able to regain attitude control after the Orbiter undocked. The core issues associated with recovering long-term attitude control using CMGs are described as well as the systems engineering analysis to identify recovery options. It is shown that the recovery method can be separated into a procedure for rate damping to a safe harbor gravity gradient stable orientation and a capability to maneuver the vehicle to the necessary initial conditions for long term attitude hold. A manual control option using Soyuz and Progress vehicle thrusters is investigated for rate damping and maneuvers. The issues with implementing such an option are presented and the key issue of closed-loop stability is addressed. A new non-propulsive alternative to thruster control, Zero Propellant Maneuver (ZPM) attitude control method is introduced and its rate damping and maneuver performance evaluated. It is shown that ZPM can meet the tight attitude and rate error tolerances needed for long term attitude control. A combination of manual thruster rate damping to a safe harbor attitude followed by a ZPM to Stage long term attitude control orientation was selected by the Anomaly Resolution Team as the alternate attitude control method for such a contingency.

  19. Summary of the 2012 Inductive Pulsed Plasma Thruster Development and Testing Program

    Science.gov (United States)

    Polzin, K. A.; Martin, A. K.; Eskridge, R. H.; Kimberlin, A. C.; Addona, B. M.; Devineni, A. P.; Dugal-Whitehead, N. R.; Hallock, A. K.

    2013-01-01

    Inductive pulsed plasma thrusters are spacecraft propulsion devices in which energy is capacitively stored and then discharged through an inductive coil. While these devices have shown promise for operation at high efficiency on a range of propellants, many technical issues remain before they can be used in flight applications. A conical theta-pinch thruster geometry was fabricated and tested to investigate potential improvements in propellant utilization relative to more common, flat-plate planar coil designs. A capacitor charging system is used to permit repetitive discharging of thrusters at multiple cycles per second, with successful testing accomplished at a repetition-rate of 5 Hz at power levels of 0.9, 1.6, and 2.5 kW. The conical theta-pinch thruster geometry was tested at cone angles of 20deg, 38deg, and 60deg, with single-pulse operation at 500 J/pulse and repetitionrate operation with the 38deg model quantified through direct thrust measurement using a hanging pendulum thrust stand. A long-lifetime valve was designed and fabricated, and initial testing was performed to measure the valve response and quantify the leak rate at beginning-of-life. Subscale design and testing of a capacitor charging system required for operation on a spacecraft is reported, providing insights into the types of components needed in the circuit topology employed. On a spacecraft, this system would accept as input a lower voltage from the spacecraft DC bus and boost the output to the high voltage required to charge the capacitors of the thruster.

  20. Electrospray Thrusters for Attitude Control of a 1-U CubeSat

    Science.gov (United States)

    Timilsina, Navin

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

  1. Low Mass Low Power Hall Thruster System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — In NASA is seeking an electric propulsion system capable of producing 20mN thrust with input power up to 1000W and specific impulse ranging from 1600-3500 seconds....

  2. Low Mass Low Power Hall Thruster System Project

    Data.gov (United States)

    National Aeronautics and Space Administration — NASA is seeking electric propulsion systems capable of producing up to 20mN thrust, input power up to 1000W and specific impulse ranging from 1600-3500 seconds. The...

  3. Magnetic Nozzle Simulation Studies for Electric Propulsion

    Science.gov (United States)

    Tarditi, Alfonso

    2010-11-01

    Electric Propulsion has recently re-gained interest as one of the key technologies to enable NASA's long-range space missions. Options are being considered also in the field of aneutronic fusion propulsion for high-power electric thrusters. To support these goals the study of the exhaust jet in a plasma thruster acquires a critical importance because the need of high-efficiency generation of thrust. A model of the plasma exhaust has been developed with the 3D magneto-fluid NIMROD code [1] to study the physics of the plasma detachment in correlation with experimentally relevant configurations. The simulations show the role of the plasma diamagnetism and of the magnetic reconnection process in the formation of a detached plasma. Furthermore, in direct fusion-propulsion concepts high-energy (MeV range) fusion products have to be efficiently converted into a slower and denser plasma jet (with specific impulse down to few 1000's seconds, for realistic missions in the Solar System). For this purpose, a two-stage conversion process is being modeled where high-energy ions are non-adiabatically injected and confined into a magnetic duct leading to the magnetic nozzle, transferring most of their energy into their gyro-motion and drifting at slower speed along with the plasma propellant. The propellant acquires then thermal energy that gets converted into the direction of thrust by the magnetic nozzle. [1] C. R. Sovinec et al., J. Comput. Phys. 195, 355 (2004).

  4. Predicting Hall Thruster Operational Lifetime Using a Kinetic Plasma Model and a Molecular Dynamics Simulation Method Project

    Data.gov (United States)

    National Aeronautics and Space Administration — Hall thrusters are being considered for many space missions because their high specific impulse delivers a larger payload mass fraction than chemical rockets. With a...

  5. The effect of easily ionized elements Na and K on the performance of pulsed plasma thruster using water propellant

    Institute of Scientific and Technical Information of China (English)

    2010-01-01

    In view of the low thrust power ratio caused by the high resistance of pulsed plasma thruster using water propellant,the paper argues that the easily ionized elements Na and K with low ionic potentials are added in the water propellant to improve its performance. The measurement of the discharging current and plasma emission spectrographic analysis prove the improvement. The experiments show that the elements Na and K have certain effect on the improvement of the performance of pulsed plasma thruster: In comparison with water propellant,the NaCl and KCl water propellant has a lower total resistance and a higher ratio of thruster power and specific impulse,and the NaCl water propellant has a slightly stronger effect on pulsed plasma thruster than the KCl. The plasma emission spectrographic analysis is in consistent with the experiment of measuring the discharging current: The elements Na and K can intensify the plasma emission spectrographic signal.

  6. Development of a cavity ring-down spectroscopy sensor for boron nitride sputter erosion in Hall thrusters

    CERN Document Server

    Tao, Lei; Gallimore, Alec D; Yalin, Azer P

    2010-01-01

    Sputter erosion of boron nitride (BN) is a critically important process in Hall thrusters from the point of view of both lifetime assessment and contamination effects. This contribution describes the development of a laser based sensor for in situ monitoring of sputtered BN from Hall thrusters. We present a continuous-wave cavity ring-down spectroscopy (cw-CRDS) system and its demonstrative measurement results from BN sputtering experiments.

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

    Science.gov (United States)

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

    2015-01-01

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

  8. The microwave thermal thruster and its application to the launch problem

    Science.gov (United States)

    Parkin, Kevin L. G.

    Nuclear thermal thrusters long ago bypassed the 50-year-old specific impulse (Isp) limitation of conventional thrusters, using nuclear powered heat exchangers in place of conventional combustion to heat a hydrogen propellant. These heat exchanger thrusters experimentally achieved an Isp of 825 seconds, but with a thrust-to-weight ratio (T/W) of less than ten they have thus far been too heavy to propel rockets into orbit. This thesis proposes a new idea to achieve both high Isp and high T/W The Microwave Thermal Thruster. This thruster covers the underside of a rocket aeroshell with a lightweight microwave absorbent heat exchange layer that may double as a re-entry heat shield. By illuminating the layer with microwaves directed from a ground-based phased array, an Isp of 700--900 seconds and T/W of 50--150 is possible using a hydrogen propellant. The single propellant simplifies vehicle design, and the high Isp increases payload fraction and structural margins. These factors combined could have a profound effect on the economics of building and reusing rockets. A laboratory-scale microwave thermal heat exchanger is constructed using a single channel in a cylindrical microwave resonant cavity, and new type of coupled electromagnetic-conduction-convection model is developed to simulate it. The resonant cavity approach to small-scale testing reveals several drawbacks, including an unexpected oscillatory behavior. Stable operation of the laboratory-scale thruster is nevertheless successful, and the simulations are consistent with the experimental results. In addition to proposing a new type of propulsion and demonstrating it, this thesis provides three other principal contributions: The first is a new perspective on the launch problem, placing it in a wider economic context. The second is a new type of ascent trajectory that significantly reduces the diameter, and hence cost, of the ground-based phased array. The third is an eclectic collection of data, techniques, and

  9. Modeling the response of ON and OFF retinal bipolar cells during electric stimulation.

    Science.gov (United States)

    Werginz, P; Benav, H; Zrenner, E; Rattay, F

    2015-06-01

    Retinal implants allowing blind people suffering from diseases like retinitis pigmentosa and macular degeneration to regain rudimentary vision are struggling with several obstacles. One of the main problems during external electric stimulation is the co-activation of the ON and OFF pathways which results in mutual impairment. In this study the response of ON and OFF cone retinal bipolar cells during extracellular electric stimulation from the subretinal space was examined. To gain deeper insight into the behavior of these cells sustained L-type and transient T-type calcium channels were integrated in the synaptic terminals of reconstructed 3D morphologies of ON and OFF cone bipolar cells. Intracellular calcium concentration in the synaptic regions of the model neurons was investigated as well since calcium influx is a crucial parameter for cell-to-cell activity between bipolar cells and retinal ganglion cells. It was shown that monophasic stimulation results in significant different calcium concentrations in the synaptic terminals of ON and OFF bipolar cells. Intracellular calcium increased to values up to fourfold higher in the OFF bipolar model neuron in comparison to the ON bipolar cell. Furthermore, geometric properties strongly influence the activation of bipolar cells. Monophasic, biphasic, single and repetitive pulses with similar lengths, amplitudes and polarities were applied to the two model neurons. Copyright © 2014 The Authors. Published by Elsevier Ltd.. All rights reserved.

  10. Cassini Spacecraft In-Flight Swap to Backup Attitude Control Thrusters

    Science.gov (United States)

    Bates, David M.

    2010-01-01

    NASA's Cassini Spacecraft, launched on October 15th, 1997 and arrived at Saturn on June 30th, 2004, is the largest and most ambitious interplanetary spacecraft in history. In order to meet the challenging attitude control and navigation requirements of the orbit profile at Saturn, Cassini is equipped with a monopropellant thruster based Reaction Control System (RCS), a bipropellant Main Engine Assembly (MEA) and a Reaction Wheel Assembly (RWA). In 2008, after 11 years of reliable service, several RCS thrusters began to show signs of end of life degradation, which led the operations team to successfully perform the swap to the backup RCS system, the details and challenges of which are described in this paper. With some modifications, it is hoped that similar techniques and design strategies could be used to benefit other spacecraft.

  11. Effect of the Hollow Cathode Heat Power on the Performance of an Hall-Effect Thruster

    Institute of Scientific and Technical Information of China (English)

    NING Zhongxi; YU Daren; LI Hong; YAN Guojun

    2009-01-01

    Effect of the hollow cathode heat power on the performance of a Hall-effect thruster is investigated. The variations in the Hall-effect thruster's performance (thrust, specific impulse and anode efficiency) with the hollow cathode heat power was obtained from the analysis of the experimental data. Through an analysis on the coupling relationship between the electrons emitted from the hollow cathode and the environmental plasma, it was found that the heat power would affect the electron emission of the emitter and the space potential of the coupling zone, which would lead to a change in the effective discharge voltage. The experimental data agree well with the results of calculation which can be used to explain the experimental phenomena.

  12. Cassini Spacecraft In-Flight Swap to Backup Attitude Control Thrusters

    Science.gov (United States)

    Bates, David M.

    2010-01-01

    NASA's Cassini Spacecraft, launched on October 15th, 1997 and arrived at Saturn on June 30th, 2004, is the largest and most ambitious interplanetary spacecraft in history. In order to meet the challenging attitude control and navigation requirements of the orbit profile at Saturn, Cassini is equipped with a monopropellant thruster based Reaction Control System (RCS), a bipropellant Main Engine Assembly (MEA) and a Reaction Wheel Assembly (RWA). In 2008, after 11 years of reliable service, several RCS thrusters began to show signs of end of life degradation, which led the operations team to successfully perform the swap to the backup RCS system, the details and challenges of which are described in this paper. With some modifications, it is hoped that similar techniques and design strategies could be used to benefit other spacecraft.

  13. Experimental identification of an azimuthal current in a magnetic nozzle of a radiofrequency plasma thruster

    Science.gov (United States)

    Takahashi, Kazunori; Chiba, Aiki; Komuro, Atsushi; Ando, Akira

    2016-10-01

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

  14. A Tool Measuring Remaining Thickness of Notched Acoustic Cavities in Primary Reaction Control Thruster NDI Standards

    Science.gov (United States)

    Sun, Yushi; Sun, Changhong; Zhu, Harry; Wincheski, Buzz

    2006-01-01

    Stress corrosion cracking in the relief radius area of a space shuttle primary reaction control thruster is an issue of concern. The current approach for monitoring of potential crack growth is nondestructive inspection (NDI) of remaining thickness (RT) to the acoustic cavities using an eddy current or remote field eddy current probe. EDM manufacturers have difficulty in providing accurate RT calibration standards. Significant error in the RT values of NDI calibration standards could lead to a mistaken judgment of cracking condition of a thruster under inspection. A tool based on eddy current principle has been developed to measure the RT at each acoustic cavity of a calibration standard in order to validate that the standard meets the sample design criteria.

  15. Space charge saturated sheath regime and electron temperature saturation in Hall thrusters

    Science.gov (United States)

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

    2005-07-01

    Existing electron-wall interaction models predict that secondary electron emission in Hall thrusters is significant and that the near-wall sheaths are space charge saturated. The experimental electron-wall collision frequency is computed using plasma parameters measured in a laboratory Hall thruster. In spite of qualitative similarities between the measured and predicted dependencies of the maximum electron temperature on the discharge voltage, the deduced electron-wall collision frequency for high discharge voltages is much lower than the theoretical value obtained for space charge saturated sheath regime, but larger than the wall recombination frequency. The observed electron temperature saturation appears to be directly associated with a decrease of the Joule heating rather than with the enhancement of the electron energy loss at the walls due to a strong secondary electron emission. Another interesting experimental result is related to the near-field plasma plume, where electron energy balance appears to be independent on the magnetic field.

  16. Iodine Hall Thruster Propellant Feed System for a CubeSat

    Science.gov (United States)

    Polzin, Kurt A.; Peeples, Steven

    2014-01-01

    The components required for an in-space iodine vapor-fed Hall effect thruster propellant management system are described. A laboratory apparatus was assembled and used to produce iodine vapor and control the flow through the application of heating to the propellant reservoir and through the adjustment of the opening in a proportional flow control valve. Changing of the reservoir temperature altered the flowrate on the timescale of minutes while adjustment of the proportional flow control valve changed the flowrate immediately without an overshoot or undershoot in flowrate with the requisite recovery time associated with thermal control systems. The flowrates tested spanned a range from 0-1.5 mg/s of iodine, which is sufficient to feed a 200-W Hall effect thruster.

  17. Study of monopropellants for electrothermal thrusters. Evaluation test program task summary report

    Science.gov (United States)

    Kuenzly, J. D.

    1974-01-01

    An electrothermal thruster designed for operation with MIL-grade hydrazine is suitable for operation with propellants having lower freezing points. These propellants are 76% hydrazine - 24% hydrazine azide, Aerozine-50, 50% hydrazine - 50% monomethylydrazine, and a TRW-formulated mixture of 35% hydrazine - 50% monomethylhydrane - 15% ammonia. A steady-state specific impulse of 200 sec was exceeded by all propellants. A pulse-mode value of 175 sec specific impulse was exceeded by the azide blend for pulse widths greater than 50 ms and was met by the carbonaceous propellants for pulse widths greater than 100 ms. Longer residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and density of screen packing. A substantial amount of thermal energy must be supplied to initiate decomposition of propellants containing unsymmetrical-dimethylhydrazine and monomethylhydrazine. The rate controlling factor appeared to be the endothermic removal of methyl radicals.

  18. Silicon Carbide (SiC) Power Processing Unit (PPU) for Hall Effect Thrusters

    Science.gov (United States)

    Reese, Bradley

    2015-01-01

    Arkansas Power Electronics International (APEI), Inc., is developing a high-efficiency, radiation-hardened 3.8-kW SiC power supply for the PPU of Hall effect thrusters. This project specifically targets the design of a PPU for the high-voltage Hall accelerator (HiVHAC) thruster, with target specifications of 80- to 160-V input, 200- to 700-V/5A output, efficiency greater than 96 percent, and peak power density in excess of 2.5 kW/kg. The PPU under development uses SiC junction field-effect transistor power switches, components that APEI, Inc., has irradiated under total ionizing dose conditions to greater than 3 MRad with little to no change in device performance.

  19. Radioisotope electric propulsion of sciencecraft to the outer solar system and near-interstellar space

    Energy Technology Data Exchange (ETDEWEB)

    Noble, R.J.

    1998-08-01

    Recent results are presented in the study of radioisotope electric propulsion as a near-term technology for sending small robotic sciencecraft to the outer Solar System and near-interstellar space. Radioisotope electric propulsion (REP) systems are low-thrust, ion propulsion units based on radioisotope electric generators and ion thrusters. Powerplant specific masses are expected to be in the range of 100 to 200 kg/kW of thrust power. Planetary rendezvous missions to Pluto, fast missions to the heliopause (100 AU) with the capability to decelerate an orbiter for an extended science program and prestellar missions to the first gravitational lens focus of the Sun (550 AU) are investigated.

  20. Ionization and Charge Exchange Reactions in Neutral Entrainment of a Field Reversed Configuration Thruster

    Science.gov (United States)

    2012-07-16

    non - Maxwellian . This indicates that a kinetic approach has to be used to model neutral entrainment in FRC thrusters. Strong impact of electron...radiative cooling can be problematic for high-Z plasma (due to a Z2 dependence) and in radiative non -equilibrium conditions (volumetric emission).1...dg, (1) where g is the relative collision velocity and fe is the Maxwellian distribution function. 2. Single charge exchange (SCX, A+ +A → A+A+) For

  1. Hall-effect thruster--Cathode coupling: The effect of cathode position and magnetic field topology

    Science.gov (United States)

    Sommerville, Jason D.

    2009-12-01

    Hall-effect thruster (HET) cathodes are responsible for the generation of the free electrons necessary to initiate and sustain the main plasma discharge and to neutralize the ion beam. The position of the cathode relative to the thruster strongly affects the efficiency of thrust generation. However, the mechanisms by which the position affects the efficiency are not well understood. This dissertation explores the effect of cathode position on HET efficiency. Magnetic field topology is shown to play an important role in the coupling between the cathode plasma and the main discharge plasma. The position of the cathode within the magnetic field affects the ion beam and the plasma properties of the near-field plume, which explains the changes in efficiency of the thruster. Several experiments were conducted which explored the changes of efficiency arising from changes in cathode coupling. In each experiment, the thrust, discharge current, and cathode coupling voltage were monitored while changes in the independent variables of cathode position, cathode mass flow and magnetic field topology were made. From the telemetry data, the efficiency of the HET thrust generation was calculated. Furthermore, several ion beam and plasma properties were measured including ion energy distribution, beam current density profile, near-field plasma potential, electron temperature, and electron density. The ion beam data show how the independent variables affected the quality of ion beam and therefore the efficiency of thrust generation. The measurements of near-field plasma properties partially explain how the changes in ion beam quality arise. The results of the experiments show that cathode position, mass flow, and field topology affect several aspects of the HET operation, especially beam divergence and voltage utilization efficiencies. Furthermore, the experiments show that magnetic field topology is important in the cathode coupling process. In particular, the magnetic field

  2. CubeSat Packaged Electrospray Thruster Evaluation for Enhanced Operationally Responsive Space Capabilities

    Science.gov (United States)

    2011-03-24

    Richard Branam, guided me through each obstacle I encountered. The AFIT laboratory staff always provided me with the tools, experience, and expertise...theorized that the thruster could be installed in the twenty degrees off vertical orientation on the thrust balance. Then using trigonometry , the...results and analysis. These sections also contain the obstacles and some solutions to the experimentation. The experimentation for the fourth

  3. Experimental Study of the Plume Characteristics of an Aged Monopropellant Hydrazine Thruster

    Science.gov (United States)

    1979-04-01

    surfaces such as solar panels, thermal control coatings , and optical surfaces can degrade satellite performance. Experimental studies with regard... coated , fused silica lens onto the thruster axial centerline. The port through which the beam passed into the chamber was also made of fused silica. As...copper alone, beeswax melted onto the copper, and SEM quality Microstick ® glue dripped onto the copper and spread by moving the disks to and fro

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

    Science.gov (United States)

    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.

  5. Development of the Multiple Use Plug Hybrid for Nanosats (MUPHyN) miniature thruster

    Science.gov (United States)

    Eilers, Shannon

    The Multiple Use Plug Hybrid for Nanosats (MUPHyN) prototype thruster incorporates solutions to several major challenges that have traditionally limited the deployment of chemical propulsion systems on small spacecraft. The MUPHyN thruster offers several features that are uniquely suited for small satellite applications. These features include 1) a non-explosive ignition system, 2) non-mechanical thrust vectoring using secondary fluid injection on an aerospike nozzle cooled with the oxidizer flow, 3) a non-toxic, chemically-stable combination of liquid and inert solid propellants, 4) a compact form factor enabled by the direct digital manufacture of the inert solid fuel grain. Hybrid rocket motors provide significant safety and reliability advantages over both solid composite and liquid propulsion systems; however, hybrid motors have found only limited use on operational vehicles due to 1) difficulty in modeling the fuel flow rate 2) poor volumetric efficiency and/or form factor 3) significantly lower fuel flow rates than solid rocket motors 4) difficulty in obtaining high combustion efficiencies. The features of the MUPHyN thruster are designed to offset and/or overcome these shortcomings. The MUPHyN motor design represents a convergence of technologies, including hybrid rocket regression rate modeling, aerospike secondary injection thrust vectoring, multiphase injector modeling, non-pyrotechnic ignition, and nitrous oxide regenerative cooling that address the traditional challenges that limit the use of hybrid rocket motors and aerospike nozzles. This synthesis of technologies is unique to the MUPHyN thruster design and no comparable work has been published in the open literature.

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

    Science.gov (United States)

    2014-06-01

    peculiar phenomenon where spokes mode is observed but only over a small section of the thruster annulus . Figures 14 and 15 show the theta-t diagrams...al., though the view was not perpendicular to the annulus plane and did not capture azimuthal plasma movement. The instrument Darnon, et. al...camera. The view of the camera is actually 4° off to the side instead of perfectly straight-on due to the need to accommodate an IR camera. This small

  7. Propeller Design Optimization for Tunnel Bow Thrusters in the Bollard Pull Condition

    Science.gov (United States)

    2012-06-01

    studies examining near wake conditions at true and near bollard pull conditions. Using a Dynamic Positioning (DP) thruster model in the Italian Ship...See the PBD User Manual (Chrisospathis, 2001 for more detail). 88 mtf2blades.in PBDOUT.MTF 0,1 blades. runloop1.bat REM Batch...OUTVEL.tec cd ..\\mtflow ..\\bin\\bl2body del ..\\prop\\veljoin.tec copy veljoin.tec ..\\prop REM Change to rotor directory and run VELCON, PBD14 and

  8. Impact of nozzle separation on the plumes of two parallel thrusters

    Science.gov (United States)

    Grabe, Martin; Dettleff, Georg; Hannemann, Klaus

    2016-11-01

    Two identical, interacting plumes emanating from model thrusters with parallel axes separated from 50 to 150 throat diameters are studied numerically. The nozzle throat Reynolds number is set to nearly 15, 000 to match that of a small bi-propellant attitude control thruster, but the simulated gas is nitrogen with a stagnation temperature of 300 K. The near-isentropic, dense plume core is computed with the DLR Navier-Stokes solver TAU and the conditions at a suitably defined interface are then used on the inflow boundary of a separately conducted direct simulation Monte Carlo (DSMC) simulation. The results are shown to agree favorably with particle flux measurements performed in the DLR high-vacuum plume test facility for chemical thrusters (STG-CT). Varying the nozzle separation distance alters the degree of rarefaction in the interaction plane, and by tagging DSMC particles according to their origin, the effect on the individual plume may be investigated. The impact of nozzle axis separation on mass flux along the line formed by the intersecting planes of symmetry is compared to the case of two equal superposed (i. e. non-interfering) plumes.

  9. Investigation on plume interference effect of solid propellant micro-thruster

    Institute of Scientific and Technical Information of China (English)

    ZHANG Bin; MAO Gen-wang; HU Song-qi; CHEN Mao-lin

    2011-01-01

    The three-dimensional numerical simulation of two-phase plume flow of solid propellant micro-thrusters was developed.Then it was used to investigate the plume interference effect by combining the direct simulation Monte Carlo(DSMC) method for multi-component gas flow with the two-way coupling model for two-phase rarefied flow.At different space between the two micro-thrusters and different wall temperature,the plume interference effect was analyzed specifically.The results show that under the plume interference effect the gas is compressed and the flow direction is changed,which resulted in the increasing of gas pressure and temperature;solid phase made no significant effect on the flow parameters of gas phase;with the rising of the space between the two micro-thrusters,the maximum pressure decreased and the maximum temperature increased in the domain under the plume interference effect;the wall temperature could influence the temperature of the gas which is extremely close to the wall,but not the gas pressure.

  10. Thrust vector control of upper stage with a gimbaled thruster during orbit transfer

    Science.gov (United States)

    Wang, Zhaohui; Jia, Yinghong; Jin, Lei; Duan, Jiajia

    2016-10-01

    In launching Multi-Satellite with One-Vehicle, the main thruster provided by the upper stage is mounted on a two-axis gimbal. During orbit transfer, the thrust vector of this gimbaled thruster (GT) should theoretically pass through the mass center of the upper stage and align with the command direction to provide orbit transfer impetus. However, it is hard to be implemented from the viewpoint of the engineering mission. The deviations of the thrust vector from the command direction would result in large velocity errors. Moreover, the deviations of the thrust vector from the upper stage mass center would produce large disturbance torques. This paper discusses the thrust vector control (TVC) of the upper stage during its orbit transfer. Firstly, the accurate nonlinear coupled kinematic and dynamic equations of the upper stage body, the two-axis gimbal and the GT are derived by taking the upper stage as a multi-body system. Then, a thrust vector control system consisting of the special attitude control of the upper stage and the gimbal rotation of the gimbaled thruster is proposed. The special attitude control defined by the desired attitude that draws the thrust vector to align with the command direction when the gimbal control makes the thrust vector passes through the upper stage mass center. Finally, the validity of the proposed method is verified through numerical simulations.

  11. Development of a Hardware-in-the-loop Simulator for Spacecraft Attitude Control Using Thrusters

    Science.gov (United States)

    Koh, Dong-Wook; Park, Sang-Young; Kim, Do-Hee; Choi, Kyu-Hong

    2009-03-01

    In this study, a Hardware-In-the-Loop (HIL) simulator using thrusters is developed to validate the spacecraft attitude system. To control the attitude of the simulator, eight cold gas thrusters are aligned with roll, pitch and yaw axis. Also linear actuators are applied to the HIL simulator for automatic mass balancing to compensate the center of mass offset from the center of rotation. The HIL simulator consists of an embedded computer (Onboard PC) for simulator system control, a wireless adapter for wireless network, a rate gyro sensor to measure 3-axis attitude of the simulator, an inclinometer to measure horizontal attitude, and a battery set to supply power for the simulator independently. For the performance test of the HIL simulator, a bang-bang controller and Pulse-Width Pulse-Frequency (PWPF) modulator are evaluated successfully. The maneuver of 68 deg. in yaw axis is tested for the comparison of the both controllers. The settling time of the bang-bang controller is faster than that of the PWPF modulator by six seconds in the experiment. The required fuel of the PWPF modulator is used as much as 51% of bang-bang controller in the experiment. Overall, the HIL simulator is appropriately developed to validate the control algorithms using thrusters.

  12. Discharge Oscillations in a Permanent Magnet Cylindrical Hall-Effect Thruster

    Science.gov (United States)

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

    2009-01-01

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

  13. VASIMR VX-200 thruster throttling optimization from 30 to 200 kW

    Science.gov (United States)

    Squire, Jared; Olsen, Chris; Chang-Diaz, Franklin; Longmier, Benjamin; Ballenger, Maxwell; Carter, Mark; Glover, Tim; McCaskill, Greg

    2012-10-01

    The VASIMR^ VX-200 experimental plasma thruster incorporates a 40 kW helicon plasma source with a 180 kW Ion Cyclotron Heating (ICH) acceleration stage integrated in a superconducting magnet. Argon propellant mass flow is injected up to 140 mg/s. Rapid plasma start up ( 10^5 liters/s) in a 150 m^3 vacuum chamber achieve performance measurements with the charge exchange mean-free-path greater than 1 m in the background neutral gas (pressure < 10-5 Torr). The thruster efficiency at 200 kW total power is 72 ± 9%, the ratio of effective jet power to input RF power, with an Isp = 4900 ± 300 seconds (flow velocity of 49 km/s), and an ion flux of 1.7 ± 0.1 x 10^21/s. The thrust increases steadily with power to 5.8 ± 0.4 N until the power is maximized and there is no indication of saturation. The plasma density near the device exit exceeds 10^18 m-3 with a power density over 5 MW/m^2. An extensive study of thruster performance, efficiency and thrust-to-power ratio, as a function of Ar propellant flow rate and ICH-to-helicon RF power ratio has been carried out over a total power range of 30 to 200 kW. Optimized throttling set points are determined. The experimental configuration and results of this study are presented.

  14. Ion properties in a Hall current thruster operating at high voltage

    Science.gov (United States)

    Garrigues, L.

    2016-04-01

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

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

    Energy Technology Data Exchange (ETDEWEB)

    Lee, B. C. [Physics Department, Colorado State University, Fort Collins, Colorado 80521 (United States); Huang, W. [NASA Glenn Research Center, 2100 Brookpark Rd., Cleveland, Ohio 44135 (United States); Tao, L.; Yamamoto, N.; Yalin, A. P., E-mail: ayalin@engr.colostate.edu [Mechanical Engineering Department, Colorado State University, Fort Collins, Colorado 80521 (United States); Gallimore, A. D. [Aerospace Engineering Department, University of Michigan, Ann Arbor, Michigan 48109 (United States)

    2014-05-15

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

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

    Science.gov (United States)

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

    2010-01-01

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

  17. Hot-Fire Testing of 5N and 22N HPGP Thrusters

    Science.gov (United States)

    Burnside, Christopher G.; Pedersen, Kevin W.; 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 hot-fire testing of 5N and 22N HPGP thrusters at the Marshall Space Flight Center’s Component Development Area altitude test stand in April 2015. Both thrusters are ground test articles and not flight ready units, but are representative of potential flight hardware with a known path towards flight application. The purpose of the 5N testing was to perform facility check-outs and generate a small set of data for comparison to ECAPS and Orbital ATK data sets. The 5N thruster performed as expected with thrust and propellant flow-rate data generated that are similar to previous testing at Orbital ATK. Immediately following the 5N testing, and using the same facility, the 22N testing was conducted on the same test stand with the purpose of demonstrating the 22N performance. The results of 22N testing indicate it performed as expected.The results of the hot-fire testing are presented in this paper and presentation.

  18. A New Method for Analyzing Near-Field Faraday Probe Data in Hall Thrusters

    Science.gov (United States)

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

    2013-01-01

    This paper presents a new method for analyzing near-field Faraday probe data obtained from Hall thrusters. Traditional methods spawned from far-field Faraday probe analysis rely on assumptions that are not applicable to near-field Faraday probe data. In particular, arbitrary choices for the point of origin and limits of integration have made interpretation of the results difficult. The new method, called iterative pathfinding, uses the evolution of the near-field plume with distance to provide feedback for determining the location of the point of origin. Although still susceptible to the choice of integration limits, this method presents a systematic approach to determining the origin point for calculating the divergence angle. The iterative pathfinding method is applied to near-field Faraday probe data taken in a previous study from the NASA-300M and NASA-457Mv2 Hall thrusters. Since these two thrusters use centrally mounted cathodes the current density associated with the cathode plume is removed before applying iterative pathfinding. A procedure is presented for removing the cathode plume. The results of the analysis are compared to far-field probe analysis results. This paper ends with checks on the validity of the new method and discussions on the implications of the results.

  19. Velocity Plume Profiles for Hall Thrusters Using Laser Diagnostic

    Science.gov (United States)

    2010-06-01

    gaseous propellant, electrical power, cooling water, USB adapters, fiber optic cables and the many instruments into the chamber. A rotary vane ...include air-cooled compressors connected to the cryo-head through flex lines to compress liquid helium and cool the cryo-heads to temperatures of 16...to the lock-in amplifiers. In addition, the housing uses a Buna 2-044 O- ring to help seal the box and a manual sliding shutter over the opening. The

  20. Neutral Entrainment Demonstration in a Xenon FRC Thruster Experiment

    Science.gov (United States)

    2013-03-01

    an ionizing collision event. The primary loss mechanism for all fundamental non -thermal electric propulsion devices and technologies is plasma...of the plasma is that they typically do not relax to Maxwellian energy distributions and so the ionization potential of the high-energy electron...Additionally, the expanded radial plasma has been greatly accelerated (compared to the non -accelerated radial profile). This expansion will have to be