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The near-term potential for iodine propellant in Hall thrusters is explored. The merits of iodine with respect to other propellants are presented. Recent performance measurements are summarized, and new measurements taken with an 8 kW thruster are presented. Thruster discharge power exceeded 10 kW, and peak measured anode efficiency exceeded 65%. Spacecraft interactions issues are also addressed and relevant data taken with a 1 kW thruster are presented. Plume data showed lower divergence with iodine than with xenon, and that a plume shield could effectively attenuate the far field plume. Material samples placed in the plume showed a strong reaction with iron, but little reaction with typical spacecraft materials. System level benefits, including low storage pressure and extremely high density, are also discussed. All results so far indicate iodine is a viable propellant for electric rockets, and for some missions is superior to xenon.

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Multiple high density propellant alternatives for Hall Effect Thrusters were investigated experimentally. The metals magnesium and zinc showed promise for high specific impulse applications. The halogen iodine showed great promise for general use, with measured performance similar to xenon and lower plume divergence. The metal bismuth showed promise for missions requiring high thrust to power. All of the high density options could substantially increase the change in velocity available to spacecraft when discharge potential and propellant volume are fixed or limited. NomenclatureB = magnetic field E = electric field e = charge of an electron, 1.6 x 10 -19 C F = thrust Φ = potential difference 0 g = gravitational constant at Earth's surface, 9.81 m/s 2 I = current, subscripts b for beam, c for cathode, d for discharge, m for magnet sp I = specific impulse j = current density m = mass, subscripts p for propellant, 0 for initial m = mass flow rate, subscript a for anode M = ion mass P = power, subscript d for discharge p = pressure, subscript s for sensor, x for xenon q = ion charge R = radial direction d V = discharge potential v = exhaust, particle, or beam velocity V ∆ = change in spacecraft velocity Z = axial direction η = efficiency, subscript t for thrust

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