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The readiness of commercial Hall thruster technology is evaluated for near-term use on competitively-award, cost-capped science missions like the NASA Discovery program. Scientists on these programs continue to place higher demands on mission performance that must trade against the cost and performance of propulsion system options. Solar electric propulsion (SEP) systems can provide enabling or enhancing capabilities to several missions, but the widespread and routine use of SEP will only be realized through aggressive cost and schedule risk reduction efforts. Significant cost and schedule risk reductions can potentially be realized with systems based on commercial Hall thruster technology. The abundance of commercial suppliers in the United States and abroad provides a sustainable base from which Hall thruster systems can be cost-effectively obtained through procurements from existing product lines. A Hall thruster propulsion system standard architecture for NASA science missions is proposed. The BPT-4000 from Aerojet is identified as a candidate for near-term use. Differences in qualification requirements between commercial and science missions are identified and a plan is presented for a low-cost, low-risk delta qualification effort. Mission analysis for Discovery-class reference missions are discussed comparing the relative cost and performance benefits of a BPT-4000 based system to an NSTAR ion thruster based system. The BPT-4000 system seems best suited to destinations located relatively close to the sun, inside approximately 2 AU. On a reference near Earth asteroid sample return mission, the BPT-4000 offers mass performance competitive with or superior to NSTAR at much lower cost. Additionally, it is found that a low-cost, mid-power commercial Hall thruster system may be a viable alternative to aerobraking for some missions.Suggestions for the near-and far-term implementation of commercial Hall thrusters on NASA science missions are discussed.

A simple analytic multistage model is presented for combined chemical-electric orbit-raising missions. Expressions for transportation rates and optimum electric specific impulse are derived for two-stage, three-stage, variable-efficiency, and tank-limited missions of up to 100 days duration. The optimum electric specific impulse is shown to depend strongly on the specific impulse of the chemical thruster. A low-thrust-trajectory optimization model is combined with launch-vehicle performance data to derive end-to-end optimized three-dimensional chemical-electric orbit-raising profiles to geostationary orbit. Optimized profiles are derived for the Sea Launch, Ariane 4, Atlas V, Delta IV, and Proton launch vehicles. Optimum electric orbit-raising starting orbits and payload mass benefits are calculated for each vehicle. The mass benefit is shown to be between 6.1 and 7.6 kg/day with two SPT-140 thrusters, or up to 680 kg for 90 days of electric orbit raising. The optimized profiles are combined with the analytic model to a create simple parametric performance model describing multiple launch vehicles. The model is a good tool for system level analysis of electric orbit-raising missions and is shown to match calculated performance to within 13%. Nomenclatureb = fitting constant c lv = effective launch-vehicle exhaust velocity, m/s c 1 = effective on board chemical thruster exhaust velocity, m/s c 2 = effective electric thruster exhaust velocity, m/s c * 2 = optimum electric thruster effective exhaust velocity, m/s d = fitting constant, m/s g = 9.81 m/s 2 m 0 = spacecraft mass, beginning of orbit raising, kg m 1 = spacecraft mass, end of chemical orbit raising, before electric orbit raising (EOR), kg m 2 = spacecraft mass, end of orbit raising (payload mass), kġ m 2 = electric propulsion mass flow rate, kg/s P = thruster input power, W T 2 = electric propulsion device thrust, N t = time, s v chem = v for all-chemical orbit raising, m/s= v for chemical portion of a C-EOR mission, m/s v 2 = v for electric portion of a C-EOR mission, m/s v 2eff = chemical v effectively replaced by electric v, m/s η p = thruster efficiency η v = mission planning efficiency

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