On-a-chip microdischarge thruster arrays inspired by photonic device technology for plasma television

This study shows that the practical scaling of a hollow cathode thruster device to MEMS level should be possible albeit with significant divergence from traditional design. The main divergence is the need to operate at discharge pressures between 1-3bar to maintain emitter diameter pressure products of similar values to conventional hollow cathode devices. Without operating at these pressures emitter cavity dimensions become prohibitively large for maintenance of the hollow cathode effect and without which discharge voltage would be in the hundreds of volts as with conventional microdischarge devices. In addition this requires sufficiently constrictive orifice diameters in the 10µm – 50µm range for single cathodes or <5µm larger arrays. Operation at this pressure results in very small Debye lengths (4 -5.2pm) and leads to large reductions in effective work function (0.3 – 0.43eV) via the Schottky effect. Consequently, simple work function lowering compounds such as lanthanum hexaboride (LaB6) can be used to reduce operating temperature without the significant manufacturing complexity of producing porous impregnated thermionic emitters as with macro scale hollow cathodes, while still operating <1200°C at the emitter surface. The literature shows that LaB6 can be deposited using a variety of standard microfabrication techniques

Microthrusters based on the T5 and T6 hollow cathodes

Small spacecraft with limited on-board resources would benefit greatly from the development of a low power, low cost microthruster able to offer propellant savings over conventional alternatives and enable higher energy missions. Such a thruster would also be beneficial in the development of all-electric spacecraft whereby the normally separate reaction control system and primary electric propulsion system were able to operate from a common propellant management system.In recent years experiments on a T6 hollow cathode have demonstrated the possibility of such a device however the performance, in particular thrust efficiency, falls far short of what would be termed a viable thruster. Nevertheless the means by which hollow cathode discharges are able to produce thrust corresponding to very high velocity propellants is not at all understood, nor is the means by which to improve performance. The relevance of the very high energy ion signatures typical of hollow cathode discharges, synonymous with hollow cathode erosion, hasalso not been established.Indirect thrust measurements were made at the University of Southampton on two separate types of hollow cathode, derived from the T5 and T6 gridded ion thrusters, with unique design modifications, primarily of anode geometry. Testing was conducted with argon and xenon and measurements were made via a deflected pendulum micro-thrust balance and supporting architecture constructed specifically for the work. Ion energy measurements were made using a hemispherical energy analyzer in a separate experiment at NASA Jet Propulsion Laboratory on a third XIPS cathode, derived from the XIPS 25cm ion thruster, with xenon and krypton.These measurements provided unique insight into the influence of terminal parameters such as discharge current, mass flowrate and cathode/anode geometry on thrust production and downstream ion energy distributions. Significant improvements in thrust efficiency have been made with the T5 cathode and in doing so, have taken hollow cathodes a significant step forwards as a viable and competitive propulsion technology.An analysis of the thrust production is made, and on this basis conclusions are drawn on the existence of electrothermal, electromagnetic and electrostatic mechanisms specific to the cathode and operating regime, as well as their roll in other associated hollow cathode phenomenon. The main conclusions of the work are presented and recommendations made for future experimental work

Numerical simulation of the insert chemistry of the hollow cathode from the deep space 1 ion engine 30,000 Hrs life test

A model for the insert chemistry developed by the authors and based on the knowledge of the BaO – CaO – Al2O3 ternary system the ELT discharge cathode insert from the Deep Space 1 life test has been simulated. The computed data show a good agreement with the experimental one; the agreement increase with the imposition of boundary conditions closer to the experimental evidence. Tungsten deposition effect have been introduced into the model using experimental data and further improving the agreement between computed and measured data. The deposition trend found suggests the possibility of a link between barium depletion and tungsten deposition

European Student Moon Orbiter solar electric propulsion subsystem architecture: an all–electric spacecraft

This paper presents the Phase A study of the Solar Electric Propulsion subsystem selected for the ESA European Student Moon Orbiter enhanced microsatellite, performed at QinetiQ under ESA funding. To minimise mass, a so-called "all electric" approach is adopted based around the re-use of the GOCE T5 gridded ion engine and the introduction of Hollow Cathode Thrusters (HCTs) for attitude control functions. Three different subsystem architectures are considered and analyzed with reference to the mass, cost, risk and level of integration between the HCTs and the T5. The favoured system architecture that best meets the various requirements adopts a shared tank and gas flow controller between the HCTs and the T5, with power being supplied from two dedicated power processing units. The possibility of reducing the propellant requirement by using an engine gimbal mechanism is also presented. The study also demonstrates how an increase in the T5 specific impulse to higher values than used on GOCE does not offer substantial system-level mass savings in this particular case