Micropropulsion for space probes of particle cosmology

Abstract

The accumulation of orbital debris poses many risks for current and future space missions. Along with an increasing number of satellite launches forecast for the coming years, it has now become mandatory for newly launched payloads to have end-of-life disposal mechanisms. Furthermore, as the market for small satellites continues to grow, there is a demand to improve the performance of these systems to allow for more complex missions to be realised. An on-board propulsion system would allow a spacecraft to perform various orbital manoeuvres, like formation flying and altitude control, which would greatly improve the operational capabilities of the satellite. Recent advancements in Microelectromechanical systems detectors and miniaturised radio frequency transmitters are facilitating a new domain for CubeSat-based astronomy research. Such in-space observatories, consisting of a constellation of CubeSats forming large distributed apertures, may allow for a rapid cost effective means of making major scientific breakthroughs. In the field of particle cosmology, these CubeSat missions can work in conjunction with ground based observatories to conduct multiwavelength observations. One of the key components for mission success is an on-board micropropulsion systemneeded for satellite maneuverability. To date, the integration of miniaturised propulsion systems on CubeSats has proven challenging, due power and mass budget restrictions. To address these issues, a proof of concept DC discharge microthruster was developed. The system utilises a coupling of propellant ionisation and acceleration mechanisms which eradicates the need for additional components, leading to reduced mass and power consumption. In this research, experimental tests were performed to characterise the system’s feasibility for use as an on-board propulsive mechanism. Preliminary experiments were conducted whereby the relation between the thruster’s electrical behaviour against variations in the electrode aperture size were explored and the thruster’s stable operating parameters were deduced. The subsequent experiments aimed to study the effect of variations in the number of electrode extraction apertures on the discharge behaviour and the magnitude of extracted ion beam current. From the measured values the system’s main performance metrics (thrust and electrical efficiency) were inferred and used to evaluate the overall feasibility of the thruster concept.