Development of a pulsed plasma thruster for nanosatellite applications

Abstract

For space missions to become more advanced and cost-effective, the manoeuvrability of spacecraft over the long term is required. Hence, both quantitative and qualitative work should be directed towards the development and optimization of electric propulsion (EP) systems that can support the mission requirements of the new space age. After comparing various propulsion systems within a survey, it was found that the electromagnetic ablative pulsed plasma thruster was best suited for this project. This is due to the simplicity, cost-effectiveness and reliability of this type of thruster. The primary objective of this work was to design, build and test a micro pulsed plasma thruster (μPPT) for utilisation in future F’SATI nanosatellite missions. Hence, a numerical model was introduced to estimate thruster performance and system parameters. The numerical model was solved using a MATLAB® application with a dedicated guided user interface (GUI). Results from the numerical simulation led to a 4.44 μF capacitor, discharging at 1500 V, while practical designs only allowed for a 6 μF capacitor, discharging at 1200V. The experimental μPPT consists of two copper electrodes separated by a solid polytetrafluoroethylene (Teflon) propellant bar of 2.5 cm2 surface area, with a discharge energy of 4.32 J. To charge the main capacitors to 1200V, an applied voltage of 9.5 V was connected to an XP power DC-to-DC converter, where manual switches were used to control the charge and discharge cycles. Due to the impracticality of manual switching on nanosatellites, a circuit proposal is presented in appendix E for future development. The experimental tests consisted of two experiments (a 40 mm electrode and a 25 mm electrode) where the voltage discharge, current discharge and thrust were measured. Testing of the thruster involved the utilisation of a vacuum chamber, a power supply, high voltage probes and an oscilloscope. To determine the thrust of the μPPT, a torsion balance is presented from the derived Cavendish experiment. From the voltage discharge measurements, it was found that for both experiments, the thruster consists of a high system resistance of approximately 126 mΩ. From the current discharge, it is shown that the plasma resistance is 53.159 mΩ and 37.04 mΩ for experiment 1 and experiment 2 respectively. Furthermore, using a self-made thrust balance, it was calculated that a maximum thrust of 3.374 mN ± 20% was achieved for a thruster discharging at 1200 V under 0.75 Torr. Tests showed that a pressure variance leads to a variance in thrust. Therefore, extrapolating between various pressure tests, while keeping the thruster parameters constant, an estimated thrust of 270 μN was theoretically achieved. Furthermore, comparing the thrust generated in experiment 1 and experiment 2 leads to the conclusion that the physical length of the thruster has minimal effect on the performance of the μPPT. Hence, a more compact design can be achieved. This thesis presents the design of a compact μPPT system for improved spacecraft manoeuvrability that is suitable for use on future F’SATI nano-satellites. It includes improved circuitry, suggested testing equipment, additional tests to fully characterise the PPT system, and recommended future work in this regard.