Up to date technology for Micro and Nano satellites include advanced attitude control and station keeping systems, especially with regard to soon scheduled missions including formation flying. Equipping even small satellites with fully operational thruster systems becomes particularly interesting with regard to the increasing focus on de-orbiting possibilities to avoid an increasing number of small satellites out of use in LEO. Miniaturized electric propulsion therefore seems to be a promising candidate for such future mission profiles.

To guarantee a safe integration, non-toxic - easy to handle - propellants are necessary, properties that make the Ablative Pulsed Plasma Thruster (PPT) become a promising candidate (Ref [1]). A PPT is a propulsion system which ionizes and then accelerates its propellant by means of electromagnetic forces to produce thrust. Due to its mechanical simplicity and its reliability, it became the first electric propulsion system ever flown in space on the Soviet Zond 2 mission which started on Nov. 30, 1964 toward Mars (Ref [2]). Since then, PPTs were used on various missions for orbital control such as drag make up and east-west stabilization (Ref [3], [4]).

The system feature high specific impulse compared to conventional chemical thrusters, mechanical simplicity, mainly due to the utilization of a solid propellant, and therefore high reliability (Ref [5]: 3f).

Schematic PPT (Ref [6]: 264)

The thrusters main components are a capacitor for power storage, the acceleration chamber consisting of propellants and electrode “rails” and an ignition system similar to a spark plug. Opposed to its fairly simple construction, the occurring processes within the PPT are highly complex: An external triggered, high current discharge ablates and then ionizes the solid propellant, in most cases Teflon^®, before both ions and electrons are accelerated by the induced magnetic fields to exit velocities beyond 30 km/sec depending on the degree of ionization. This way the energy stored in the capacitor is depleted within micro seconds, which makes the thrusters operate in a short pulsed mode, similar and well compatible to digital logic control (Ref [5]: 4). In addition to the electromagnetic acceleration, a gas dynamic process similar to the one in an arc jet accelerates existing neutral atoms. Both processes add up to impulse bits on the range of a few tens of micro Newtons which makes the system well fitted for very small satellites, despite its significantly poor performance with efficiencies below 10%.

Trying to implement such technology into satellites in the size of Cubesats of course brings a variety of new challenges due to its restrictive mass, volume and power budget, especially regarding the energy storage unit.

The testing of such a PPT, designed within the master thesis of a SEDSAT II member at the Austrian Research Institute in Seibersdorf (A), could be a good way to draw attention of the scientific community to our project for using technology with promising importance in the near future of small satellite missions.

References:

[1] Pottinger S. J., Scharlemann C. A. (2007): Micro Pulsed Plasma Thruster Development, IEPC-2007-125.

[2] Burton, R. L., Turchi, P. J. (1998): Pulsed Plasma Thruster. Journal of Propulsion and Power, Vol 14, No 5, p 716-734.

[3] Vondra R., Thomassen K., Solbes A. (1971): A Pulsed Electric Thruster for Satellite Control, IEEE 01450062 Vol. 59. No. 2.

[4] Ziemer J. K., Choueiri E. Y. (2001): Scaling laws for electromagnetic pulsed plasma thrusters, Plasma Sources Sci. Technol. 10. p. 395-405.

[5] Guman, W. J. (1968): Pulsed Plasma Technology in Microthrusters, Fairchild Hiller Corp., Technical Report AFAPL-TR-68-132.

[6] Jahn, R. G. (1968): Physics of Electric Propulsion, Dover Publications, Inc, Mineola, New York.