Angelo Grubisic

Development of an indirect counterbalanced pendulum optical-lever thrust balance for micro- to millinewton thrust measurement

This paper describes the design and testing of an indirect hanging pendulum thrust balance using a laser-optical-lever principle to provide micro- to millinewton thrust measurement for the development of electric propulsion systems. The design philosophy allows the selection of the total thrust range in order to maximize resolution through a counterbalanced pendulum principle, as well as passive magnetic damping in order to allow relatively rapid transient thrust measurement. The balance was designed for the purpose of hollow cathode microthruster characterization, but could be applied to other electric propulsion devices in the thrust range of micro- to millinewtons. An initial thrust characterization of the T5 hollow cathode is presented.

Hollow cathode thrusters for all-electric spacecraft

This paper characterizes thrust performance of a variant of the T6 hollow cathode modified for improved operational characteristics. The device displays promising performance and highlights the potential for application as a microthruster. The T6 cathode is able to produce specific impulse of over 1050s with xenon and argon; however the dominant acceleration mechanism at low flow rates is shown to be almost entirely electromagnetic effectively constituting a low power, magneto-plasma-dynamic thruster. The modified cathode also displayed very low discharge voltages ~12V and stable operation while carrying currents as high as 30-Amps on flow rates of less than 0.1mgs-1 down to 0.04mgs-1 with xenon. This investigation also highlights the importance of electrode geometry in hollow cathode thruster design and indeed in all other applications of hollow cathode where flow rates should be minimized

Electric Propulsion Subsystem Architecture for an All-Electric Spacecraft

For many space missions, both a main propulsion subsystem and additional attitude control (AOCS) propulsion subsystem are required. These subsystems normally use different propellants, hence require separate tanks, different flow control units (FCU) and, in case of solar electric propulsion (SEP), separate power processing units (PPU). This leads to increases in total mass of the spacecraft and complexity while reducing system specific impulse. One possibility to alleviate this problem would be to develop a main and an AOCS propulsion technology which could be integrated, sharing some of the components required for their operation, hence reducing system mass. A spacecraft employing such combined technologies as part of an SEP system is referred to as an “All-electric-spacecraft” (Wells et al., 2006). In this chapter, the system design for an all-electric-spacecraft will be presented. A gridded ion engine (GIE) is proposed as a main propulsion subsystem with hollow cathode thrusters (HCT) considered for the AOCS propulsion subsystem. The mission considered during this study is the ESA European Student Moon Orbiter (ESMO), which the University of Southampton proposed to use SEP for both attitude control and main propulsion. During the ESMO phase-A study, a full design of the SEP subsystem was performed at QinetiQ as part of a wider study of the mission performed in conjunction with QinetiQ staff and funded by ESA. The output of this study will be here presented to explain the concept of the all-electric-spacecraft, its benefits, drawbacks and challenges.

Ion Energies in an Open-Diode Hollow Cathode Discharge

Ion energy measurements from a ¼ inch cathode were made with krypton and xenon expellants at a range of flow rates and discharge currents using an electrostatic energy analyzer mounted on axis. High energy ion content scaled strongly with discharge current and inversely with mass flow rate and molecular mass suggesting a possible MDH production mechanism. At high current conditions for krypton and xenon peak ion energies exceeded 130eV (>17kms-1) and 75eV (>10kms-1) with high energy tails extending well beyond 140eV and 200eV respectively. No conclusive correlation was found with discharge noise in the form of high frequency voltage oscillations and high energy ion content

Hollow Cathodes as a Plasma Propulsion Device

Testing of T5 and T6 ion thruster derived hollow cathode thrusters (HCT) at the University of Southampton has demonstrated impressive propulsive performance showing potential for a number of new applications. Ongoing studies also indicate that the hollow cathode thruster concept can be scaled down to a micro-electro-mechanical-systems (MEMs) level device with improving thrust efficiencies. Initial results indicate 3 forms of HCT propulsion device are possible. The first is a high power 1200s with inert propellants such as xenon and argon; essentially a low current (<30A) MPD-HCT thruster. The second form of device operates in an electrothermal mode at lower powers and currents (<100W, <3.2A) and is capable of delivering at least 427s with argon. More recently HCTs have also been proposed in a third form as an on-chip MEMs device for applications at very low power (<5W) with theoretical specific impulse up to 180s for xenon. This paper describes the current findings in HCT technology research and the proposed the possibility of extending this to MEMs scale

A Novel Interferometric Microwave Radiometer Concept Using Satellite Formation Flight for Geostationary Atmospheric Sounding

For most Earth observation applications, passive microwave radiometry from the geostationary orbit requires prohibitively large apertures for conventional single-satellite platforms. This paper proposes a novel interferometric technique capable of synthesizing these apertures using satellite formation flight. The significance of such concept is in its capacity to synthesize microwave apertures of conceptually unconstrained size in space for the first time. The technique is implemented in two formation flight configurations: a formation of a single full-sized satellite with microsatellites and a formation of several full-sized satellites. Practical advantages and challenges of these configurations are explored by applying them to geostationary atmospheric sounding at 53 GHz, the lowest sounding frequency considered for future sounder concepts Geostationary Atmospheric Sounder, GeoSTAR, and Geostationary Interferometric Microwave Sounder. The two configurations produce apertures of 14.4 and 28.8 m, respectively, and a spatial resolution of 16.7 and 8.3 km, respectively, from the geostationary orbit. The performance of these interferometers is simulated, and the challenges identified are threefold. First, intersatellite ranging in micrometer-level precision is required. Second, the extremely sparse design suggests that further innovation is necessary to improve radiometric resolution. Third, the presence of long baselines suggests extreme decorrelation effects are expected. While the first requirement has already been demonstrated on ground, the other two remain for future research. This technique can be implemented at arbitrary microwave frequencies and arbitrary circular orbits, meaning it can also be applied to other geostationary applications, or to achieve unprecedented spatial resolution from lower orbits, or to extend the accessible frequencies into lower frequency radio waves.

Geostationary atmospheric sounding 1 by formation flight aperture synthesis

This paper introduces a multi-satellite approach to passive microwave interferometric radiometry applied to geostationary atmospheric sounding at 53 GHz. The concept applies satellite formation flight to the currently operational interferometric techniques to extend the achievable microwave aperture sizes, leading to unprecedented spatial resolution for microwave radiometers. The presented configurations are capable of synthesising 14.4 m aperture and larger using SMOS-sized satellites, and spatial resolution better than 16.7 km at 53 GHz from the geostationary orbit. Two instrument concepts are proposed: Single-Element Companion concept and Array Duplicate concept, both of which are scalable, where the achievable spatial resolution is extendible and constrained only by the number of satellites in the constellation. The performance of both concepts are simulated, and the results show that the interferometer is highly sensitive to uncertainties in inter-satellite position measurements, where measurement errors as small as 6 mm result in 6.3 K root mean square error in radiometric accuracy.

Thrust production mechanisms in Hollow cathode microthrusters

Hollow cathode have recently been investigated at the University of Southampton as potential standalone microthrusters. 12Thrust measurements suggest that in some cases, hollow cathodes are able to generate specific impulse of over 1000s with xenon. The means by which hollow cathodes are able to generate such high levels of specific impulse is not clearly understood. This paper explores thrust production mechanisms in the T5, T6 and XIPS hollow cathodes based on thrust performance and ion energy measurements. Analysis suggests the total thrust produced includes components of gas dynamic thrust particularly as a result of an intense electron pressure at the cathode exit, but also shows evidence for magneto-hydro-dynamic (MHD) forces at high currents and low flow rates arising from the self-induced azimuthal magnetic field within the orifice and resulting cross-field interaction plasma. Data may also suggest ion acceleration due to plasma potential hills. While this initial characterization can only loosely attribute the relative magnitude each mechanism plays it does show evidence for each. This research shows that with further development based on an understanding of the thrust mechanism, hollow cathodes thrusters could present significant refinements to the technology of electric propulsion.