Testing of the Variable Specific Impulse Magnetoplasma Rocket VX-200 engine was performed over a wide throttle range in a 150m3 vacuum chamber with sufficient pumping to permit exhaust plume measurements at argon background pressures less than 1x10-3Pa (1x10-5torr) during firings, ensuring charge-exchange mean free paths longer than the vacuum chamber. Measurements of plasma flux, radio frequency power, propellant flow rate, and ion kinetic energy were used to determine the ionization cost of argon and krypton in the helicon discharge. Experimental data on ionization cost, ion fraction, exhaust plume expansion angle, thruster efficiency, and thrust are presented that characterize the VX-200 engine performance over a throttling range from 15 to 200kW radio frequency power. A semiempirical model of the thruster efficiency as a function of specific impulse indicates an ion cyclotron heating efficiency of 85 +/- 7%. Operation at a total radio frequency coupled power level of 200kW yields a thruster efficiency of 72 +/- 6% at a specific impulse of 4900 +/- 300s with argon propellant. A high thrust-to-power operating mode was characterized over a wide parameter space with a maximum thrust-to-power ratio of 51 +/- 5mN/kW at a specific impulse of 1660 +/- 100s for a ratio of ion cyclotron heating radio frequency power to helicon radio frequency power of 0.7:1.

Testing of the VX-200 engine was performed over a wide throttle range in a vacuum facility with sufficient volume and pumping to permit exhaust plume measurements at low background pressures and sufficiently large charge exchange mean free paths. Experimental results are presented with the VX-200 engine installed in a 150 m 3 vacuum chamber with an operating pressure below 1x10 -3 Pa (1x10 -5 Torr), and with an exhaust plume diagnostic measurement range of 5 m in the axial direction and 1 m in the radial directions. Measurements of plasma flux, RF power, and neutral gas flow rate, combined with knowledge of the kinetic energy of the ions leaving the VX-200 engine, are used to determine the ionization cost of argon and krypton in the first stage helicon plasma discharges. Experimental data on ionization cost, ion fraction, exhaust plume expansion angle, thruster efficiency, and total force are presented that characterize the VX-200 engine performance over a wide throttling range from 15 kW to 200 kW total RF power. A semi-empirical model of the thruster efficiency as a function of specific impulse was developed to fit the experimental data, and reveals a second stage Ion Cyclotron Heating (ICH) RF power coupling efficiency of 85%. Operation at an RF power level of 200 kW yields a thruster efficiency of 72%±9% at a specific impulse of 4900±300 s. A high thrust-to-power operating mode was characterized over a wide parameter space with a maximum thrust to power ratio of 51 mN/kW at a specific impulse of 1660 s for a ratio of ICH RF power to helicon RF power of 0.7:1.

IGH-POWER electric propulsion thrusters can reducepropellant mass for heavy-payload orbit-raising missions andcargo missions to the moon and near-Earth asteroids, and they canreduce the trip time of robotic and piloted planetary missions [1–4].TheVariableSpecificImpulseMagnetoplasmaRocket(VASIMR®)VX-200 engine is an electric propulsion system capable ofprocessing power densities on the order of 6MW

The helicon plasma stage in the Variable Specific Impulse Magnetoplasma Rocket (VASIMR (R)) VX-200i device was used to characterize an axial plasma potential profile within an expanding magnetic nozzle region of the laboratory based device. The ion acceleration mechanism is identified as an ambipolar electric field produced by an electron pressure gradient, resulting in a local axial ion speed of Mach 4 downstream of the magnetic nozzle. A 20 eV argon ion kinetic energy was measured in the helicon source, which had a peak magnetic field strength of 0.17 T. The helicon plasma source was operated with 25 mg s(-1) argon propellant and 30 kW of RF power. The maximum measured values of plasma density and electron temperature within the exhaust plume were 1 x 10(20) m(-3) and 9 eV, respectively. The measured plasma density is nearly an order of magnitude larger than previously reported steady-state helicon plasma sources. The exhaust plume also exhibits a 95% to 100% ionization fraction. The size scale and spatial location of the plasma potential structure in the expanding magnetic nozzle region appear to follow the size scale and spatial location of the expanding magnetic field. The thickness of the potential structure was found to be 10(4) to 10(5) lambda(De) depending on the local electron temperature in the magnetic nozzle, many orders of magnitude larger than typical laboratory double layer structures. The background plasma density and neutral argon pressure were 10(15) m(-3) and 2 x 10(-5) Torr, respectively, in a 150 m(3) vacuum chamber during operation of the helicon plasma source. The agreement between the measured plasma potential and plasma potential that was calculated from an ambipolar ion acceleration analysis over the bulk of the axial distance where the potential drop was located is a strong confirmation of the ambipolar acceleration process.

® ) VX-200, a 200 kW flight-technology prototype. Results are presented from first stage only and first stage with booster stage experiments that were performed on the VX-200 using between 60 mg/s and 150 mg/s argon propellant. Measurements of ion flux, ion energy, plasma density and potential gradients, and force density profiles taken in the exhaust plume of the VX-200 are made within a 150 cubic meter vacuum chamber and are presented in the context of individual stage and total engine performance. An emphasis will be given to the current status of technology innovation that makes the ground based VX-200 and the ISS mounted VF-200 engines possible.

The VASIMR[R] Flight Experiment (VF-200-1) will be tested in space aboard the International Space Station (ISS) in about four years. It will consist of two 100 kW parallel plasma engines with opposite magnetic dipoles, resulting in a near zero-torque magnetic system. Electrical energy will come from ISS at low power level, be stored in batteries and used to fire the engine at 200 kW. The VF-200-1 project will provide a unique opportunity on the ISS National Laboratory for astrophysicists and space physicists to study the dynamic evolution of an expanding and reconnecting plasma loop. Here, we review the status of the project and discuss our current plans for computational modeling and in situ observation of a dynamic plasma loop on an experimental platform in low-Earth orbit. The VF-200-1 project is still in the early stages of development and we welcome new collaborators.

A 200 kW VASIMR R (cid:13) engine designed specifically to demonstrate the end-to-end DC electrical power conversion to thrust power has undergone initial testing in a new vacuum facility and successfully operated with a low temperature superconducting magnet, two solid-state RF generators, and improved RF coupler designs. This engine is given the name VX-200 for a VASIMR R (cid:13) experimental device that operates at 200 kW input electrical power. The VX-200 has shown through its successful operation that the primary technologies required for a spaceflight version of the VASIMR R (cid:13) are compatible with each other and can operate at practical efficiencies. The first stage, or helicon section, of the engine has operated at full power with maximum magnetic field. The second stage, or booster, has operated at nearly full power with a lower magnetic field. Ion flux measurements were taken in a new 150 m 3 vacuum chamber with 100,000 liters/second of pumping that contained the VX-200. The first stage generated an argon plasma jet with a cost to extract an electron-ion pair of 78 ± 11 eV/ion at 32 kW and a flow rate of 135 mg/s. A record power of 149.2 kW was coupled to the plasma with the booster RF coupler.

(Abstract) The Variable Specific Impulse Magnetoplasma Rocket (VASIMR™) is a high power magnetoplasma rocket, capable of Isp/thrust modulation at constant power. The plasma is produced by a helicon discharge. The bulk of the energy is added by ion cyclotron resonance heating (ICRH.) Axial momentum is obtained by adiabatic expansion of the plasma in a magnetic nozzle. Thrust/specific impulse ratio control in the VASIMR™ is primarily achieved by the partitioning of the RF power to the helicon and ICRH systems, with the proper adjustment of the propellant flow. Ion dynamics in the exhaust were studied using probes, gridded energy analyzers (RPA's), microwave interferometry and optical techniques. This paper will summarize results from high power ICRH experiments performed on the VX-100 using argon plasma during 2007. An overview of the way forward will be touched on briefly, with some emphasis on the fact that VASIMR™ is now being developed by private enterprise. The opportunities and challenges of this situation will be reviewed. VASIMR™ was originally designed to serve as the sustainer engines for the manned Mars mission, for which it remains one of the leading candidate systems. A number of other uses of the VASIMR™ have been studied in some detail, including robotic missions to the outer planets, ISSO reboost, station keeping and sustained maneuvering of robotic craft in Earth orbit, and lunar cargo hauling. We will explore the latter in this paper. A number of studies have illustrated the cost and mass efficiency of solar-electric propulsion as an alternative to chemical propulsion for hauling cargo from low Earth orbit to low lunar orbit; recent studies considered the Hall thruster in this application. Here, we present the results of a payload vs. specific impulse trade study for a six-month Earth-Moon transit time, and compare the technical and economic features of the VASIMR™ technology to other electric thrusters for this application.

An observed 20 eV argon ion energy is attributed to a measured axial plasma potential profile within the expanding magnetic nozzle region of the Variable Specific Impulse Magnetoplasma Rocket (VASIMR ® ) VX-200i device, a 10% field version of the VX-200 prototype. The ion acceleration mechanism is identified as an ambipolar flow caused by expanding plasma that follows an idealized electron Boltzmann relation, resulting in a maximum axial speed of S c 1 . 4 ~ . The VX-200i prototype was operated with 25 mg/s argon