Electrospray Propulsion Systems for Small Satellites

For this paper, Busek presents its heuristic for evaluating capability of propulsion systems for spacecraft at the CubeSat level, which utilizes the well-known rocket equation, but recognizes that for very small spacecraft the mass ratio (initial mass/final mass) has much greater influence than the specific impulse. This provides a simple basis of comparison for various propulsion technologies, as well as prioritization of development objectives for improving overall baseline capabilities. Examples of selected propulsion technologies are presented for illustration, followed by an introduction of Busek’s different propulsion solutions for Cubist-scale spacecraft. Finally, Busek discusses its proposed electrospray thrusters for the DARPA Phoenix mission, with a brief overview of the Phoenix objectives and how the rocket equation heuristic, plus additional system considerations, led to selection of electrospray propulsion as the optimal mission solution. The importance of propulsion in maximizing spacecraft capability cannot be underestimated, providing both mobility and maneuverability that would otherwise be limited or nonexistent. While there are a variety of challenges in implementing propulsion in small spacecraft, such as scaling down of thrusters or operating in a limited power scenario, perhaps the greatest challenge is of physics itself, where the low available mass and volume necessarily result in spacecraft with low mass ratios, where the mass ratio is mi/mf, (initial mass/final mass), and mi-mf = mp, the mass of propellant. Figure 1 illustrates qualitatively the relationship , (go=9.8m/s 2 ) showing the self-evident dependence of ΔV upon both the Isp of the propulsion system and the amount of propellant carried relative to the overall mass of the spacecraft. Less evident, however, is how low the mass ratio actually is for spacecraft as the spacecraft decreases in size, and the consequent ceiling in available ΔV. A typical 3U/3kg CubeSat, in order to maximize payload and accommodate its other required components, often is unable to sacrifice more than 1⁄2 U, or 500cm 3 , for a propulsion system. Assuming that this translates to 0.5kg propellant, this 3U CubeSat carrying 1⁄2 kg of propellant with Isp=200s achieves ΔV≈ 350m/s. In reality, however, given the required hardware comprising a propulsion system, propellant loading, and commensurate ΔV, is often much lower, usually <50m/s. Figure 1 Illustration of effect of mass ratio (initial mass/final mass) and Isp upon ΔV This gives rise to the question of how best to evaluate propulsion system candidates for miniature spacecraft. By subjecting the propulsion system to very stringent mass and volume constraints, and with the dry mass of small propulsion systems comprising a significant amount of this mass and volume budget, high Isp systems ought to be of particular interest. After further