Dr Aaron Knoll

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%.

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 Halo thruster is a cusped field thruster currently under investigation at the Surrey Space Centre, University of Surrey. The concept concerns the addition of a novel toroidal cusp layer, the ?halo?, to a magnetic field topology closely resembling that of the Cylindrical Hall Thruster. Preliminary low resolution maps are presented of plasma potential, electron temperature and plasma density within an electromagnet Halo thruster discharge channel, produced using a 2-axis translating Langmuir probe arrangement. Potential drops are observed in both the annular and cylindrical parts of the discharge channel, and a region of high plasma density is revealed on the central axis in the cylindrical part, suggesting performance may be improved by adopting a purely cylindrical geometry.

A plasma thruster comprises a plasma chamber having first and second axial ends, the first of which is open, an anode located at the second axial end, and a cathode. The cathode and anode are arranged to produce an electric field having at least a component in the axial direction of the thruster. A magnet system comprising a plurality of magnets is spaced around the thruster axis, each magnet having its north and south poles spaced around the axis.

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.

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.

A simple model is presented for the time-averaged electron mobility within a Hall thruster. The model is predicated on the experimental evidence for isentropic electron flow and, when used in a one-dimensional simulation, captures plasma properties that are in reasonable agreement with experiment.

The Canadian Space Agency's (CSA) Software and Ground Segment Section (SGS) has the mandate to develop innovative emerging software and on-board satellite and ground segment computer technologies. To that end, there is an ongoing development of a simulation environment to test COTS (Commercial-Of-The-Shelf) technologies. There are severe cost constraints in all aspects of many space missions due to the limited return on investment and scarce commercialization opportunities that come with many science missions. There is an opportunity to explore the innovative implementation of COTS technologies to reduce the mission cost and maximize performance available from COTS components. However, using COTS technologies in the space environment has ist constraints and therefore designing a spacecraft mission has to involve some new techniques that allow implementation of these components and minimize the risk of failure. The goal of our project is to develop a simulation environment, itself using COTS components, and then to allow the seamless integration of various components to test spacecraft mission concepts. For example, one of the aspects of using COTS processors in space is to protect them from the radiation environment. The current state of the simulation tests an innovative software EDAC (Error Detection and Correction) package and a redundant processor configuration to investigate protection against the effects of radiation and other failures on a generic mission. It also includes the capability to test formation-flying concepts that have the potential to revolutionize cost reduction efforts for space missions and to enable new space applications. This paper describes the simulation environment in detail and illustrates some of the technologies being tested for possible future space missions. The paper concludes with a look at the future development of the simulation environment and possible benefits of its use as well as the lessons learned to date.

The Canadian Space Agency?s (CSA) Software and Ground Segment (SGS) section has the mandate to develop innovative software and ground segment technologies. The implementation of formation flying concepts for Canadian missions is also currently under investigation at CSA's Spacecraft Engineering section. To that end, there is an ongoing development of a simulation environment to test COTS (Commercial-Off-The-Shelf) and formation flying technologies. Some of today's spacecraft are laboriously custom designed for a specific mission and a limited set of tasks. Development time can be lengthy (several years), which means that designs do not take advantage of the most recent technology. Designs also tend to be extremely inflexible, creating a spacecraft that cannot be easily adapted to future missions. A design ethic that promotes reusability is much more cost-effective and increases the time available to advance new technologies. COTS offer advantages such as a reduced development time, an increased product selection, faster and cheaper parts replacement, and extensively tested advanced designs. The main drawbacks to COTS use in space are susceptibility to radiation and in some cases decreased reliability. Since one of the main advantages of formation flying is the reduced mission sensitivity to a spacecraft failure, the risk associated with COTS, which has hindered its use in conventional space missions, is less concerning in the context of a multiple spacecraft mission. Achieving some level of standardization is a problem currently confronting the space industry, which must be addressed to realize the cost savings that can come from mass production and spacecraft interoperability. The use of standard components with standard interfaces also reduces development time. As well, since part of the goal is to have spacecraft already in orbit regroup and possibly join with new spacecraft to accomplish other missions, some forwards and backwards compatibility between generations of spacecraft will be necessary. This paper describes an integrated simulation environment that uses COTS spacecraft and simulation components to investigate formation flying scenarios and their benefits and challenges. A few of the well-known industry software and hardware tools incorporated into the environment include Analytical Graphic's STK, Mathwork's Matlab/Simulink, CAE Electronic's Real-time Object-Oriented Simulation Environment (ROSETM), Intel's StrongARM processor, and the PC-104

This paper presents results on the development of both hybrid and multifluid simulations of Hall thrusters that resolve azimuthal electron flow dynamics. Simulations are carried out for a laboratory, nominally 90 mm channel diameter discharge with an extended acceleration region for which a modest collection of experimental data exists. The simulations are intended as a tool to better understand the mechanism behind azimuthal wave-driven electron transport. Both the hybrid and fluid simulations capture azimuthal fluctuations which appear to be consistent with quasi-neutral disturbances predicted by linear analysis. The impact of such disturbances on the cross-field transport is discussed.

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.

A rescue device includes a control module to sense an avalanche, sense the direction of the surface and establish a target path to the surface. A nozzle is selected or oriented by the control module along the target path. A fluid reservoir is connected to the nozzle to force a fluid through the nozzle along the target path to the surface. This allows rescuers to identify the location of a victim and also provides an air path to the victim.

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.

This paper describes a 2-dimensional simulation of a coaxial Hall thruster that was developed in the axial-azimuthal (z - ¸) 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.

An ion accelerator includes: an inner magnet having a channel extending through it in an axial direction; an outer magnet extending around the inner magnet, the magnets having like polarities so as to produce a magnetic field having two locations of zero magnetic field strength. The locations are spaced apart in the axial direction; and an anode and a cathode are arranged to generate an electrical potential difference between the locations.

A 3-axis translating Langmuir probe system has been developed in order to characterize the plasma properties in the plume region of the Quad Confinement Thruster. These experiments have provided a volumetric map of the electron density, plasma potential, and electron temperature for three operating conditions of the device. The operating points investigated in this study were for a symmetric magnetic field and two magnetic steering scenarios where alternate sets of the thruster?s four electromagnets were powered. The most significant outcome of these experiments was a quantitative analysis of the thrust vectoring capabilities of the device showing an 11 degree angular offset of the plasma beam from the thruster axis for each of the two steering scenarios investigated. It was also found that there was a 90 degree change in orientation between the two steering conditions. This study has also contributed insight into the 3-dimensional structure of the plasma plume. The results indicate that the electron density and plasma potential peak near the centre axis whereas the electron temperature is at a minimum.

The Hall thruster is a type of plasma propulsion system for space vehicle applications. The thrust produced by this device is derived from the momentum of ions, which are accelerated to high exit velocities by the action of an electric field sustained within the plasma. The advantage of the Hall thruster compared to conventional chemical rocket propulsion is a significantly higher exhaust velocity, which leads to better utilization of propellant mass. Since the early days of Hall thruster research, experiments have suggested that the mobility of electrons along the axis of the thruster, perpendicular to an imposed magnetic field, is higher than can be explained by classical collision transfer processes alone. A lack of understanding regarding the mechanism for this enhanced mobility has proved a significant challenge toward the development of reliable simulations capable of predicting the performance of these devices. This thesis examines the role of high frequency plasma oscillations on the electron mobility using a combination of experimental studies on a laboratory Hall thruster, and numerical simulations capable of capturing these oscillations and quantifying their impact on the electron mobility. Two high frequency oscillations were consistently observed in the experiments: a 10MHz mode which appeared strongest in the vicinity of the anode, and a 4.5MHz mode which was strongest in the mid-channel region of the thruster. These were relatively low wave number (long wavelength) oscillations: approximately 6cm for the 4.5MHz oscillation and 3cm for the 10MHz oscillation. The angle of these waves varied considerably depending on the operating conditions of the thruster. They were found to be closely aligned to the axis of the thruster for experiments conducted with Xenon propellant, and were aligned with the circumference of the thruster (in the direction of electron drift) for experiments conducted with Krypton. A Hall thruster simulation, formulated in the axial-azimuthal coordinates of the thruster, was able to capture high frequency oscillations in reasonable agreement with experimental findings: 13MHz near the anode and 5MHz in the mid-channel region of the thruster for 160V discharge conditions. The simulation results demonstrated the crucial role of these oscillations in regulating the electron transport. In the vicinity of these oscillations the electron mobility was increased by a factor of five or more. The central finding of this thesis is that

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 multi-fluid 2-D simulation of a co-axial E x B plasma discharge is presented, resolving the azimuthal dynamics leading to the growth and saturation of high-frequency (0.5 -- 10 MHz) azimuthally-propagating fluctuations. The simulation accounts for finite-rate ionization kinetics, with associated losses of particles and energy to the bounding ceramic walls. These discharges are typical of Hall thruster plasma accelerators, which are increasingly being used in space propulsion applications. The simulations presented are for full scale thrusters that operate in the 1 kW power levels, capturing the entire azimuthal domain. The simulations focus on the role played by these fluctuations in establishing the cross-field electron current in regions of relatively strong magnetic fields (50-200 Gauss). The time-average predictions for plasma properties are in qualitative and quantitative agreement with experiments, and the findings seem to be supportive of the experimental results that indicate that high frequency fluctuations may be more important at defining electron current at lower discharge voltages, where the azimuthal electron shear is small.

A new variety of electric thruster, the Quad Confinement Thruster, has been designed and tested during a preliminary one year development effort. This thruster utilizes an innovative magnetic cusp topology. Eight electromagnets are used to create a convex magnetic field structure with a center cusp, and four outlying cusps along the periphery. The thrust produced by the device is derived from the momentum of ions accelerated through a Hall effect static electric field. Direct thrust measurements show a specific impulse up to 700s at powers of less than 100W. This specific impulse was achieved using a propellant flow rate of 5sccm of Krypton, and corresponded to a thrust of approximately 2.1mN. One of the key motivations for this research is the ability of this thruster to actively control the direction of thrust through manipulation of the magnetic field by regulating power to the individual solenoid magnets. Preliminary experiments, which measure the profile of the plasma density at two axial stations downstream of the channel exit, appear to show a 14 degree thrust vectoring capability of the device.

A plasma torch having an open end from which a plasma plume is emitted in use is disclosed. The plasma torch includes a central cathode rod, a grounded conductive tube having an open end and being arranged around the cathode and spaced therefrom to form a first cylindrical cavity open at one end; and a high voltage electrode having a dielectric barrier material at a radially inward-facing surface thereof and being arranged around the grounded conductive tube and spaced apart therefrom to form a second annular cylindrical cavity open at one end. A constant direct current (DC) electrical power plus a high voltage pulsed electrical power is provided to the cathode producing an arc discharge in the first cavity between the cathode and grounded tube to generate a central thermal plasma emitted at an open end of the first cylindrical cavity. A high voltage alternating current electrical power or pulsed electrical power is provided to the high voltage electrode producing a dielectric barrier discharge in the second annular cylindrical cavity to generate a non-thermal plasma emitted from an open end of the second cavity as a halo around the central thermal plasma.

We present an inter-laboratory comparison of the performance and plasma plume measurements of the 200W Quad Confinement Thruster (QCT-200) between the Surrey Space Centre (SSC) electric propulsion laboratory and the ESA Propulsion Laboratory (EPL) at ESA-ESTEC. The test campaign involves thrust balance measurements of the QCT-200 device over a range of operating conditions, and plasma plume measurements using Faraday probes. A matching set of test conditions following a common test procedure is conducted in both facilities and the results critically compared.

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.

We report on progress towards the development of a Hall thruster simulation in the axial-azimuthal (z - ¸) 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 - ¸ 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.

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%.

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.

This study examines the detailed dispersion properties of high frequency (1- 15 MHz) plasma fluctuations within a Hall thruster. The results reveal the existence of two strong modes in this frequency range, which are predominantly axially propagating, with weaker azimuthal waves near the exit and near field region. Both modes show generally similar dispersion characteristics. Each propagates predominantly towards the anode within the discharge, and towards the cathode in the near field. Each is characterized by low wavenumber cut-offs (near 4.5 MHz and 10.5 MHz respectively). In all cases, within the channel, the results seem to be only weakly dependent on axial position, indicating that these disturbances are highly non-local. The wavenumber lies within a range of 100-700 rad/m, and the phase velocities fall in the range of 105 5x105 m/s. The azimuthal components near the exit of the discharge have phase velocities approaching the azimuthal electron drift velocity, but their dispersive characteristics fall short of what is expected for a typical beam plasma mode. We suspect that the because the frequency of these waves lie close to the lower-hybrid resonance, that these waves are the result of an interaction between the lower hybrid waves and the electron stream mode, driven and coupled by the relatively strong transverse electron shear flow, the scale length of which corresponds closely to the measured wavelengths.

Electric Propulsion (EP) systems can enable novel spacecraft missions requiring high total change in velocity, owing to their high speci?c impulse compared to chemical propulsion systems. Mature devices, such as Hall E?ect Thrusters (HETs), have accumulated signi?cant ?ight heritage. How- ever, established technologies do not satisfy the requirements of the rapidly growing small satellite sector, because of adverse scaling to low powers. The Halo thruster concept falls within the cat- egory of Cusped Field Thrusters (CFTs), aimed at addressing this issue. The concept concerns the use of ?magnetic null regions?, formed through the deliberate cancellation of magnetic ?elds. Two such regions are produced in the thruster, a ?null point? at the thruster exit and an annular ?halo? near the anode. The work presented in this thesis has provided foundational knowledge of the performance and internal physics of the Halo thruster, using a 5 cm channel diameter, electromagnet laboratory model. Measurements of thrust, speci?c impulse and thrust e?ciency were obtained over a wide range of operating conditions using a pendulum thrust balance in representative high vacuum, and the sensitivity of the measured performance to facility e?ects was assessed. Trends in plasma potential, electron temperature and plasma density internal to the discharge channel were obtained using a translating Langmuir probe, allowing the basic physics of operation of the device to be inferred. The thruster was found to exhibit comparable performance to other CFTs, with measurements shown to be robust to facility e?ects. Internal plasma measurements revealed behaviour similar to that of the Cylindrical Hall Thruster, with some di?erences due to the presence of the halo magnetic null region near the anode which might be exploited to improve performance. As a result of the research presented, design changes are suggested for future iterations. In its current embodiment, the thruster already o?ers advantages over heritage small satellite EP systems, and is a viable candidate for near-term industrialisation.

The design and performance of a novel direct current (dc) neutralizer for electric propulsion applications are presented. The neutralizer exploits an E × B discharge to enhance ionization via electron-neutral collisions. Tests are performed with helium, argon, xenon, air, and water vapor as working gases. The I-V characteristics and extraction parameters are measured for both atomic and molecular gases. The maximum partial power efficiency is 4.2 mA/W in argon, 2.7 mA/W in air, and 2 mA/W in water vapor. The typical utilization factor is below 1 and the power consumption is less than 120 W. A semiempirical model is derived to predict the performance of dc plasma cathodes using atomic gas. A comparison with existing plasma cathodes and conventional LaB6 cathodes is presented, and design optimizations aimed at improving the performance are proposed.

We characterize the ion velocity flow field in the plasma ejected from a Quad Confinement Thruster using non-intrusive 2-D laser-induced fluorescence diagnostics. Measurements show a free-space ion acceleration layer located 8 cm downstream of the exit plane, with an observed ion velocity increase from 3 km/s to 10 km/s within a region of 1 cm thickness or less. The ion velocity field is investigated with different magnetic configurations, demonstrating how distorting the magnetic field produces changes in ion velocity magnitude and direction as well as in metastable (probed) ion density.

In this paper we present the design and test campaign results of two plasma cathodes for electric propulsion applications. One cathode is based on a Hall-type discharge operated in DC. Three magnetic topologies have been tested in order to govern the discharge and the electron extraction with this neutralizer. The second cathode exploits a planar magnetron discharge operated in DC. Preliminary results of extraction tests involving atomic and molecular gases are presented. It is shown that the presence of open-loop Hall currents and null magnetic regions created by four arc magnets whose axial polarity is alternated may increase extraction performance of the Hall-type neutralizer. It is also shown that the extraction characteristics of the planar magnetron neutralizer are qualitatively similar to those of the Hall-type neutralizer.

The first flight unit of the 200 W class Quad Confinement Thruster will be demonstrated in orbit on the SSTL NovaSAR spacecraft. Key preparatory activities have involved extensive ground testing in order to identify the operational and performance envelopes of the thruster over a broad range of test conditions with the ultimate aim of accurately predicting the in-space behavior. In particular, experimental campaigns have been carried out at the Surrey Space Centre and ESA Propulsion Laboratory at ESA-ESTEC in the effort to determine vacuum facility effects on the measured parameters through a critical comparison of the results obtained in the different laboratories.