Technical Notes VX-200 Magnetoplasma Thruster Performance Results Exceeding Fifty-Percent Thruster Efficiency Benjamin W. Longmier, ∗ Leonard D. Cassady, † Maxwell G. Ballenger, ‡ Mark D. Carter, § Franklin R. Chang-Díaz, ¶ Tim W. Glover, ∗∗ Andrew V. Ilin, †† Greg E. McCaskill, ‡‡ Chris S. Olsen, §§ and Jared P. Squire ¶¶ Ad Astra Rocket Company, Webster, Texas 77598 and Edgar A. Bering III ∗∗∗ University of Houston, Houston, Texas 77204 DOI: 10.2514/1.B34085 I. Introduction H IGH-POWER electric propulsion thrusters can reduce propellant mass for heavy-payload orbit-raising missions and cargo missions to the moon and near-Earth asteroids, and they can reduce the trip time of robotic and piloted planetary missions [1–4]. The Variable Specific Impulse Magnetoplasma Rocket (VASIMR®) VX-200 engine is an electric propulsion system capable of processing power densities on the order of 6 MW=m 2 with a high specific impulse (4000 to 6000 s) and an inherent capability to vary the thrust and specific impulse at a constant power. The potential for a long lifetime is due primarily to the radial magnetic confinement of both ions and electrons in a quasi-neutral flowing plasma stream, which acts to significantly reduce the plasma impingement on the walls of the rocket core. High-temperature ceramic plasma-facing surfaces handle the thermal radiation: the principal heat transfer mechanism from the discharge. The rocket uses a helicon plasma source [5,6] for efficient plasma production in the first stage. This plasma is energized further by an ion cyclotron heating (ICH) RF stage that uses left-hand polarized slow-mode waves launched from the high field side of the ion cyclotron resonance. Useful thrust is produced as the plasma accelerates in an expanding magnetic field: a process described by conservation of the first adiabatic invariant as the magnetic field strength decreases in the exhaust region of the VASIMR [7–9]. End-to-end testing of the VX-200 engine has been undertaken with an optimum magnetic field and in a vacuum facility with suffi- cient volume and pumping to permit exhaust plume measurements at low background pressures. Experimental results are presented with the VX-200 engine installed in a 150 m 3 vacuum chamber with an operating pressure below 1  10 2 Pa (1  10 4 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 argon 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 the argon plasma. A plasma momentum flux sensor (PMFS) measures the force density as a function of radial and axial positions in the exhaust plume. New experimental data on ionization cost, exhaust plume expansion angle, thruster efficiency, and total force are presented that charac- terize the VX-200 engine performance above 100 kW. A semi- empirical model of the thruster efficiency as a function of specific impulse has been developed to fit the experimental data, and an extrapolation to 200 kW dc input power yields a thruster efficiency of 61% at a specific impulse of 4800 s. II. Experimental Setup and Method A. VX-200 Engine The VX-200 engine is an experimental VASIMR prototype designed to operate at 200 kW of input dc electrical power. The device provides an end-to-end integrated test of the primary VASIMR components in a vacuum environment, with the goal of measuring and improving the system performance. A majority of the VX-200 engine components are located within the vacuum chamber, with only the solid-state RF generators, superconducting magnet power supplies, and cryocoolers kept at atmospheric pressure. The superconducting magnet, structural, rocket core, engine sensors, and electrical components are operated within the vacuum chamber. Figure 1 shows a schematic of the VX-200 engine installed inside the vacuum chamber and the approximate shape of the magnetic field flux lines within the core and magnetic nozzle. The core of the VX-200 engine is defined as the components that physically surround the plasma and intercept the bulk of the waste heat radiated by the plasma column. The VX-200 engine is restricted to pulses of less than 1 min, owing to temperature limitations of certain seals and joints in the rocket core. The helicon stage launches a right-handed circularly polarized wave, which produces a cold argon plasma. A pressure gradient drives the plasma flow through a magnetic choke into the ICH stage, where another RF coupler launches a wave that preferentially heats the ions in a single pass [9]. The VX-200 engine RF generators convert facility dc power to RF power and perform impedance matching between the RF generator output and the rocket core. The RF generators were custom built by Nautel, Ltd., model numbers VX200-1 (helicon generator) and VX200-2 (ICH generator). The VX200-1 RF generator is rated up to 48  1 kW RF, with a 91  1% efficiency and a specific mass of 0:85  0:02 kg=kW. The VX200-2 generator is rated up to 172  1 kW RF, with a 98  1% efficiency and a specific mass of 0:506  0:003 kg=kW. The generator efficiencies were determined by independent testing performed by Nautel, Ltd., which included a direct measurement of input power and calorimetry of the dissipated power in the generator. The exhaust velocity of the ions increases as the coupled ICH power increases. Coupled RF power is defined as the RF power that is injected by the helicon and/or ICH couplers and is inductively absorbed by the plasma column or radiatively lost by the RF Received 12 August 2010; revision received 1 February 2011; accepted for publication 24 February 2011. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/11 and $10.00 in correspondence with the CCC. ∗ Principal Research Scientist; ben.longmier@adastrarocket.com. Member AIAA. † Lead Engineer; lcassady@adastrarocket.com. Member AIAA. ‡ Staff Scientist; maxwell.ballenger@adastrarocket.com. § Director of Technology; mark.carter@adastrarocket.com. Member AIAA. ¶ Chief Executive Officer;, aarc@adastrarocket.com. Associate Fellow AIAA. ∗∗ Director of Development; tim.glover@adastrarocket.com. Member AIAA. †† Computational Research Lead; andrew.ilin@adastrarocket.com. ‡‡ Senior RF Engineer; greg.mccaskill@adastrarocket.com. §§ Research Scientist; chris.olsen@adastrarocket.com. ¶¶ Director of Research; jared.squire@adastrarocket.com. Member AIAA. ∗∗∗ Professor, Departments of Physics and Electrical and Computer Engineering; eabering@uh.edu. Associate Fellow AIAA. JOURNAL OF PROPULSION AND POWER Vol. 27, No. 4, July–August 2011 915