Aaron's research interests involve experimental and computational plasma physics related to electric propulsion for spacecraft applications. His Ph.D. research focused on Hall Thrusters, and was aimed at understanding the high frequency plasma oscillations that occur in these devices and their relevance to the efficiency and performance. He also has personal research interests in Fluid/Aerodynamic simulations and computational techniques for solution of non-linear partial differential equations.
EADS, Astrium, Surrey Satellite Technology Ltd
Find me on campus Room: 01 BA 00
9:00 - 17:30, Monday - Thursday
9:00 - 17:00, Friday
Performance measurements have been obtained of a novel propulsion concept called the Halo thruster under development within the University of Surrey. The Halo thruster, a type of cusped-field thruster with close similarity to the cylindrical Hall thruster, is motivated by the need for low-power and low-cost electric propulsion for the small satellite sector. Two versions of the device are investigated in this study: a design using permanent magnets at high magnetic-field strength and a design using electromagnets with moderate field strength. While operating at 200 W discharge power, which is of particular interest to power-limited small satellite platforms, the permanent-magnet design achieved a maximum thrust efficiency of 8% at a specific impulse of approximately 900 s using a krypton propellant. By comparison, the electromagnet design achieved a maximum thrust efficiency of 28% at a specific impulse of approximately 1500 s at 200 W using a xenon propellant. For higher levels of power (tested up to 800 W), the performance of the electromagnetic design saturated at approximately 25% thrust efficiency using krypton and 30% using xenon. The thrust efficiency of the permanent-magnet design appeared to increase monotonically up to 600 W reaching a maximum value of 14%.
A thrust balance characterization of a low powered Quad Confinement Thruster is presented for high levels of propellant flow. The nominal flow rate for this device is between 1 and 2 sccm of xenon propellant. This paper extends the operating range, and investigates the performance at two high flow conditions of 10 and 20 sccm. Power is varied incrementally between 20 and 200 W in order to characterize the performance versus power trends of the device. It was found that for these high flow regimes the propellant is underutilized, and a proportion of the increased thrust can likely be attributed to a hot gas expansion of the neutral xenon rather than the generation of additional accelerated ions. The thrust was increased from 1 (nominal) to 3.3 mN at 200 W of input power for the 20 sccm condition. However, the performance penalty in terms of the specific impulse was considerable. The specific impulse under these conditions dropped below 200 s, where the nominal condition is 1000 s. A compromise between increased thrust and decreased performance was found at 10 sccm of flow: 3 mN of thrust at 300 s specific impulse.
In order to evaluate the accuracy and sensitivity of a pendulum-type thrust measurement system, a linear variable differential transformer (LVDT) and a laser optical displacement sensor have been used simultaneously to determine the displacement resulting from an applied thrust. The LVDT sensor uses an analog interface, whereas the laser sensor uses a digital interface to communicate the displacement readings to the data acquisition equipment. The data collected by both sensors show good agreement for static mass calibrations and validation with a cold gas thruster. However, the data obtained using the LVDT deviate significantly from that of the laser sensor when operating two varieties of plasma thrusters: a radio frequency (RF) driven plasma thruster, and a DC powered plasma thruster. Results establish that even with appropriate shielding and signal filtering the LVDT sensor is subject to plasma noise and radio frequency interactions which result in anomalous thrust readings. Experimental data show that the thrust determined using the LVDT system in a direct current plasma environment and a RF discharge is approximately a factor of three higher than the thrust values obtained using a laser sensor system for the operating conditions investigated. These findings are of significance to the electric propulsion community as LVDT sensors are often utilized in thrust measurement systems and accurate thrust measurement and the reproducibility of thrust data is key to analyzing thruster performance. Methods are proposed to evaluate system susceptibility to plasma noise and an effective filtering scheme presented for DC discharges.
The performance of a novel neutralizer for space applications based on a ExB discharge is presented. Preliminary tests were carried out with argon gas and flow rates in the range of 5-10 SCCM. Electrons were extracted through an orifice of diameter 1.8 mm. The maximum extracted current versus input power reported was 2.4 mA/W. The total power input, given by the sum of discharge power plus the extraction power, was in the range of 40-90 W. During extraction tests, the discharge current was limited at 0.2 A due to limit in the cooling system. Future work will be focused on tests at various extraction orifice diameters and cathode materials. Ultimately, xenon and non-conventional gases would be tested as working gases.
The effect of fuel to oxidiser ratio on the thrust performance of a novel CubeSat propulsion system is presented in this paper. This propulsion system uses aluminium wool as fuel and a mixture of water and sodium hydroxide as oxidiser. The goal of the experiment is to determine the effect of fuel to oxidiser ratio on the thrust profile of the device, as measured with a pendulum type thrust balance in a vacuum chamber facility. Experimental results show that a low fuel to oxidiser ratio reduces the propulsion efficiency and does not support multiple injections. A peak thrust value of 0.032 N was recorded with a specific impulse of 45 s. Based on this specific impulse the anticipated delta-V for a 1U CubeSat of 1.33 kg is 80 m/s, assuming a dry mass ratio of 83.33%.
There is currently a gap in the market for a low cost Electric Propulsion solution for small spacecraft. The 200W Quad Confinement Thruster (QCT-200) has the potential to fill this gap, allowing small satellites to become much more capable in terms of propulsion whilst maintaining a price point which is acceptable to customers. The device fits well with the current SSTL philosophy, and the existing SSTL xenon feed system can be simply adapted to allow the QCT-200 to effectively be a bolt on module. Surrey Satellite Technology Ltd. (SSTL), the Surrey Space Centre (SSC), and Airbus Defense and Space (AD&S) have been working on a flight standard design of the device. This paper discusses the short history of the QCT-200, the operational principle of the device, and the industrialisation of the device from its experimental origins. Finally the application of the QCT-200 in to a current spacecraft for in-orbit performance demonstration in 2016 is then related along with the mission scenarios enabled by the device.
A thrust balance characterization of a low powered Quad Confinement Thruster is presented for high levels of propellant flow. The nominal flow rate for this device is between 1sccm and 2sccm of Xenon propellant. This study extends the operating range, and investigates the performance at two high flow conditions of 10sccm and 20sccm. Power is varied incrementally between 20W and 200W in order to characterize the performance versus power trends of the device. It was found that for these high flow regimes the propellant is underutilized, and a proportion of the increased thrust can likely be attributed to a hot gas expansion of the neutral Xenon rather than the generation of additional accelerated ions. The thrust was increased from 1mN (nominal) to 3.3mN at 200W of input power for the 20sccm condition. However, the performance penalty in terms of the specific impulse was considerable. The specific impulse under these conditions dropped below 200s, where the nominal condition is 1000s. A compromise between increased thrust and decreased performance was found at 10sccm of flow: 3mN of thrust at 300s specific impulse.
The Quad Confinement Thruster employs a convex magnetic field bounded by four cusps to weakly con fine electrons and thus create a high density plasma. An electric field sustained between a rear anode and an external hollow cathode provides ion acceleration. In this study the first performance measurements of a permanent magnet high powered QCT (QCT1500) are reported. Direct thrust measurements were made, using a pendulum type thrust balance, as a function of the anode power up to maximum power of 800 W. A symmetric quadrupole field strength of 950 G was used throughout and the krypton propellant flow was varied from 10-30 sccm. Thrust levels between 3-10 mN at specific impulses of 200-1600s were recorded.
A 2-dimensional Hall thruster simulation has been developed in the axial-azimuthal coordinate plane. The goal of this simulation is to numerically model high frequency plasma waves within the discharge channel of the Hall thruster, and study the contribution of these waves to the time-averaged axial electron drift. This model uses a continuum (fluid) representation for both the electrons and ions. In order to simulate oscillations in the electron field it was necessary to model the electrons dynamically, as opposed to assuming a steady state solution at each time step. The electron momentum equations also include electron inertia terms that are normally neglected in typical Hall thruster models. These inertia terms provide a wave coupling mechanism between axially and azimuthally propagating waves. This numerical model was able to reproduce two dominant high frequency plasma oscillations in the Hall thruster: a 74MHz Kelvin-Helmholtz type shearing instability, and a 7MHz oscillation in the plasma density that has also been observed experimentally. The simulation was successful at predicting the axial electron drift in good agreement with experiment. The results of this study suggest that the plasma oscillations play a dominant role in the electron transport process. In particular, contributions to the electron transport resulting from perturbations in the azimuthal electron velocity were found to be greater than 300% of classical collisional transport.
We report on progress towards the development of a Hall thruster simulation in the axial-azimuthal (z - θ) computational space. Unlike most computational studies of closed-drift Hall accelerators which have been in one dimension (1D) along the axial direction or in two dimensions (2D) in the axial and radial dimensions, and which require some specification of the axial transport mechanism, this z - θ numerical simulation developed here self-consistently evolves the azimuthal electron drift velocity. The simulation is, in principal, capable of capturing correlated azimuthal disturbances in plasma properties which may give rise to cross-field transport, and makes no use of ad-hoc transport models. Preliminary analysis of the results indicates that azimuthal plasma instabilities may contribute to the axial electron transport process.
An experimental setup has been developed to measure high frequency plasma oscillations within the acceleration channel of a laboratory Hall thruster. The plasma oscillations are measured with three Langmuir probes separated by small axial and azimuthal offsets. This configuration permits the oscillations to be correlated with direction and wave number. This work is motivated by the anomalous electron transport phenomena, as plasma instabilities may play a crucial role in this transport process. Preliminary data has been gathered downstream of the exit plane of the thruster and suggests high frequency oscillations in the 1 to 10MHz range predominately in the axial direction. Work is currently underway to measure the high frequency oscillations within the acceleration channel at various axial locations.
This paper describes a 2-dimensional simulation of a coaxial Hall thruster that was developed in the axial-azimuthal (z - θ) computational space. Most computational studies of closed-drift Hall accelerators have been in one dimension (1D) along the axial direction or in two dimensions (2D) in the axial and radial dimensions. These 1D and 2D models have had reasonable success in describing the overall behavior of the plasma discharge. However, in these descriptions, the axial transport of electrons is modeled in an ad hoc fashion, usually with a prescribed cross-field mobility. The cross-field electron mobility is likely to be influenced/established by the azimuthal dynamics. Azimuthal perturbations arise from the established equilibrium and, if properly correlated, result in a net axial transport of electrons. The numerical model developed in this study self-consistently evolves the azimuthal electron drifts, and makes no use of ad hoc transport models. Preliminary analysis of the results indicates that azimuthal plasma instabilities do contribute to the axial electron transport process. However, both numerical and theoretical challenges still need to be addressed as there were notable discrepancies in terms of the time averaged ion velocity and electron density characteristics as compared with experimental findings. These differences are partly attributed to spurious spikes in the plasma potential, the origins of which are yet to be identified.
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