NASA required a rocket thruster able to produce a number of pulses at high specific impulse at a relatively low voltage (~300 to 400V). The key problem was that existing propellants for liquid-fueled pulsed plasma thrusters (LPPTs) required high voltages to ablate and accelerate the propellant. The Green Electric Monopropellant (GEM)-fueled Pulsed Plasma Thruster (PPT) is a rocket thruster that is able to produce a number of pulses at high specific impulse at a relatively low voltage (~300 to 400V).

The liquid GEM-type fuel has never before been considered for pulsed plasma acceleration but liquid-fueled PPTs have been investigated using water, mercury, dimethyl ether, and other propellants. The disadvantage of the prior art is the higher voltage required to ablate and accelerate the fluid.

During functional operation of the GEM LPPT, a valve dispenses a small amount of propellant between the two electrodes. Then once the valve is closed, a capacitor, or bank of capacitors, discharges across the propellant, heating and accelerating it outward to produce thrust. The advantage of the innovation is that with lower voltage, less power processing equipment is needed, and the electrodes will experience less erosion for a longer lifetime.

Test data acquired using a pendulum thrust stand was utilized to verify the performance of the LPPT concept. The motion of the pendulum was measured using a laser range finder while a scale was used to determine mass loss to permit the calculation of Isp values. Sources of error include these test hardware sensors but they only affect the actual performance data and not the goal to demonstrate a LPPT fueled by GEM.

Peripheral equipment includes the electrical power delivery system. Operating the thruster at varying discharge voltage and energy per pulse allows the mapping of the performance parameter space, which in turn aids in determining the optimum specific impulse for the design. The laboratory thruster requires the replacement of the electrodes due to erosion, but a flight system should be designed with sufficient electrode lifetime to permit the completion of its mission with only one set of electrodes. Reliability is still being determined at this time, and electrode erosion margin factors are dependent on the mission profile and the power level utilized on the LPPT.

This work was done by Jason Thrasher, Shae Williams, and Phillip Takahashi of Digital Solid State Propulsion for Marshall Space Flight Center. NASA is seeking partners to further develop this technology through joint cooperative research and development. For more information about this technology and to explore opportunities, please contact Ronald C. Darty at This email address is being protected from spambots. You need JavaScript enabled to view it.. MFS-33328-1


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This article first appeared in the March, 2019 issue of Tech Briefs Magazine.

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