Dr Andrea Lucca Fabris

An Electron Cyclotron Resonance (ECR) magnetic nozzle plasma thruster typically consists of a microwave antenna and a magnetic nozzle. The magnetic field forms a region where the propellant is resonantly heated and ionised by the microwaves. The ionised gas is then accelerated out of the magnetic nozzle, generating thrust. The design of these thrusters to date has not accounted for the thickness of the resonance region, often assuming it to be near zero. Two test campaigns have been conducted in which the thickness of the resonance region was varied. Larger resonance region thicknesses are seen to increase thrust and specific impulse by up to 60 % and thrust efficiency by up to 32 %. Experimental measurements of electron temperature, plasma potential and ion current indicate that increased electron heating in the larger resonance volume leads to a stronger electrostatic field and higher ion beam current. This study presents strong evidence that optimising for a thickened resonance region can significantly enhance thruster performance.

Spacecraft in Very Low Earth Orbit (VLEO) could play a significant role in improving daily life through enhanced communication, Earth observation and environmental monitoring. Their close proximity to Earth enables higher resolution imagery for improved weather forecasting and disaster monitoring, faster data transmission with lower latency, and faster natural deorbiting to ensure compliance with space debris mitigation policies. However, operating at such altitudes poses significant challenges, particularly due to atmospheric drag, and requires active propulsion systems to maintain operational orbits over extended periods. Additionally, Earth's gravity field variations exert greater influence on VLEO satellites. In this research, a high-fidelity orbital propagator is being developed including atmospheric drag effects calculated via simplified Gas Surface Interaction (GSI) models and perturbations from the Earth's non-spherical gravity field and third-body attractions from the Moon and the Sun. Aerodynamic forces acting on the satellite were computed using panellised satellite models, with relative velocity including horizontal neutral winds. The propagator's accuracy was verified using Gravity Field and Steady-State Ocean Circulation Explorer (GOCE) data with GSI parameter's described in the literature, achieving close alignment with GOCE's ephemeris during its free-fall. An air-breathing satellite platform aligned with the orbital velocity vector was tested with the propagator. This alignment, caused deviations in the relative particle velocity vector, resulting in elevated drag. This rendered passive compression systems insufficient to the meet required number density thresholds, necessitating the need for active intake systems. A control methodology was derived from the rate of change of perigee using Gauss Variational Equations which demonstrated improved flexibility and robustness compared to previous methods available in the literature.

A miniaturised External Discharge Plasma Thruster (XPT) is developed and characterised from 50 - 250 W. The thruster performance and discharge envelope are recorded as a function of xenon and krypton mass flow rate at various anode powers. The thrust of the miniaturised XPT on krypton decreases by up to 35 % and the anode efficiency decreases by up to 60% for a similar anode operation. The ion current density and utilisation efficiencies are estimated as the thruster operation is varied between xenon and krypton. As the mini-XPT is operated on krypton the beam current remains relatively constant as the anode power increases. At higher anode voltages the current utilisation for krypton is noted to decrease. The plasma properties are presented as the anode power is varied between xenon and krypton at a constant mass flow rate. The plasma potential and density decrease by up to 35 % as the operation is varied between 0.6 mg/s of xenon and krypton. Spatial maps are presented identifying the plasma potential and density at a constant thruster operation of 65 W and 0.6 mg/s of xenon mass flow rate. The axial electric field distribution is presented as a function of the plasma potential and axial distance away from the thruster.

The study of the solar corona has important ramifications on the understanding and forecasting of space weather phenomena. Yet, regardless of scientific breakthroughs brought by space-based coronagraphs, access to the lowest layers of the Sun’s atmosphere remains possible mostly during rare and sporadic total solar eclipses on Earth. This paper introduces the preliminary trajectory design analyses of a Moon-Enabled Sun Occultation Mission (MESOM), which capitalizes on synodic resonant orbits in the Sun-Earth-Moon four body problem to enable global and high-quality measurements of the solar corona below 1.02 sun radii once every synodic month (e.g., 29.6 days)

The study of the solar corona has important ramifications on the understanding and forecasting of coronal mass ejections, solar flares, and solar energetic particle events that can pose a significant threat to society. Yet, regardless of scientific breakthroughs brought by space-based coronagraphs, access to the lowest layers of the Sun's atmosphere remains challenging because of vignetting and stray light effects that significantly degrade signal-to-noise ratios in these regions. An alternative approach, first proposed by Eckersley and Kemble, advocates creating artificial total eclipses in space by flying a spacecraft in the shadow of the Moon. This paper introduces the preliminary trajectory design analyses and trade-off studies of a Moon-Enabled Sun Occultation Mission (MESOM). By means of synodic resonant orbits that exists in the chaotic dynamics of the Sun-Earth-Moon system, trajectories capable of delivering on average 15 minutes per synodic month (29.6 days circa) of manoeuvre-free solar corona observations below 1.02 sun radii were identified and used as a baseline for the preliminary design of a 2+ year-long satellite mission.

The quad confinement plasma source is a novel plasma device developed for space propulsion applications, whose core is an 𝐸×𝐵 discharge with open electron drift. The magnetic field is produced by independently powered electromagnets able to generate different magnetic field topologies with the ultimate aim of manipulating the ion flow field for achieving thrust vectoring. In this work, we map the ion velocity in the plasma ejected from the quad confinement thruster with different magnetic configurations using non-intrusive laser-induced fluorescence diagnostics. Measurements show a steep ion acceleration layer located 8 cm downstream the exit plane of the discharge channel, detached from any physical boundary of the plasma source. In this location, the ion velocity increases from 3 to 10 km/s within a 1 cm axial region. The ion acceleration profile has been characterized under multiple testing conditions in order to identify the influence of the magnetic field intensity and topology on this peculiar ion acceleration layer.

A torsional thrust balance has been designed and validated by Surrey Space Centre and Added Value Solutions UK Ltd. in collaboration with the UK Space Agency. The thrust stand has been tested with two electric propulsion (EP) systems operating with xenon: the Halo thruster and the XJET thruster. The first consists of a low-power (< 1 kW) Hall effect-based thruster, whose thrust level is between 3 and 20 mN, depending on the power of the system. The second is an electron cyclotron resonance thruster whose operative point is in the 0.3-1.5 mN thrust range. The thruster is mounted on a titanium rotating beam, whose movement is measured by an optical fiber displacement sensor. The thrusters' direct current electrical connections are routed through room temperature liquid metal pots and microwave power is transmitted via a wireless transfer system, minimizing friction effects. To reduce thermal issues during long thruster operations, the torsional thrust balance is designed with a water-cooling hub around the flex pivot. Noise from the laboratory environment is lessened by using four vibration-dampening spring systems as thrust balance feet. The tests on the two EP systems have shown accurate and repeatable results, demonstrating that the balance can be used to characterize different EP systems in the mu N-mN thrust range. (C) 2022 Author(s).

A hollow cathode with a modular design has been developed to assist with laboratory testing of plasmabased thrusters for satellite applications. This novel modular design includes interchangeable components for varying the geometry and tailoring the configuration to specific applications, as well as easing the replacement of individual components in the case of damage. The modular hollow cathode also presents unconventional design features aimed at improving the heating efficiency: the heater is in direct contact with the emitter and the keeper is not in physical contact with the cathode base. The modular hollow cathode development has been based on a combination of theoretical modelling and experimental testing. The influence of the novel mechanical assembly has been investigated by characterising the operational envelope at different propellant mass flow rates for xenon and krypton. The modular hollow cathode has demonstrated stable operation by sustaining discharge currents between 0.5 and 4 A at different conditions. Finally, the cathode has been coupled with a Hall-type plasma thruster operating in the 0.3–2.5 A anode current range. This paper outlines development, experimental validation of this peculiar mechanical cathode configuration, covering plasma and thermal modelling, standalone testing, and coupled Hall-type thruster operation.

PLATOR is a new electrothermal thruster for space logistics applications, developed by the University of Surrey and the University of Leicester. This paper describes the technology behind the development of the thruster and presents a mission scenario where a PLATOR-propelled spacecraft is used to capture and de-orbit the European Space Agency (ESA)'s Envisat satellite. The orbital transfer trajectory is designed using a time-optimal control approach, and the spacecraft's state vector's uncertainties are assessed through a covariance analysis. A navigation analysis is then performed to evaluate the spacecraft's capability to autonomously track its motion during the transfer using GPS measurements. Finally, a target proximity phase is then simulated to demonstrate the spacecraft's capability to rendezvous and dock with Envisat, using the uncertainties obtained from the covariance analysis, showing the potential of the PLATOR thruster for in-orbit servicing and active debris removal applications.

The behaviour of plasma within the discharge channel of the Quad Confinement Thruster is studied on the basis of electron kinetics. Here we propose that E × B drift of electrons drives the formation of unusual quadrant dependent light emitting structures observed experimentally in the discharge channel of the Quad Confinement Thruster. This assertion is made on the basis of a theory-based analysis and a computational model of the Quad Confinement Thruster. A particle orbit model of electron motion under the influence of applied electric and magnetic fields was used to assess electron transport. Structures strongly resembling that of the observed visible emission regions were found in the electron density distribution within the channel. While the motion of electrons cannot be decoupled from the motion of ions, as in this simple electron kinetic approximation, the results of this analysis strongly indicate the physical mechanism governing the formation of the non-uniform density distributions within the Quad Confinement Thruster channel.

The possibility of efficiently exploiting Very Low Earth orbits (VLEO) poses significant technological challenges. One of the most demanding constraints is the need to counteract the drag generated by the interaction of the spacecraft with the surrounding atmosphere. Funded by the European Commission under the H2020 programme, the Air-breathing Electric THrustER (AETHER) project aims at developing the first propulsion system able to maintain a spacecraft at very-low altitudes for an extended time. The main objective of the project is to demonstrate, in a relevant environment, the critical functions of an air-breathing electric propulsion system, and its effectiveness in compensating atmospheric drag. This achievement will involve multiple research activities, among which: (i) the characterization of specific application cases through an extensive market analysis in order to define specific requirements and constraints at different design levels, (ii) fulfilment of pertinent testing conditions of flight conditions on-ground, relevant to the specific mission cases, (iii) the development of critical technologies, in particular those relevant to the collection, the ionization and the acceleration of rarefied atmospheric mixtures and (iv) the testing of the RAM-EP thruster to assess the system performance. In this paper, the main activities foreseen in the AETHER project are described, providing the detailed perspective towards an effective exploitation of the project outcomes for a possible future in-orbit demonstration.

Abstract The air-breathing electric propulsion (ABEP) concept refers to a spacecraft in very-low Earth orbit (VLEO) ingesting upper atmospheric air as propellant for an electric thruster. This compensates atmospheric drag and allows the spacecraft to maintain its orbital altitude, removing the need for on-board propellant storage and allowing an extended mission duration which is not limited by propellant exhaustion. There is a need for development of a robust, high current density and long life cathode (or neutralizer) for air-breathing electrostatic thrusters as conventional thermionic hollow cathodes are susceptible to oxygen poisoning. An Air-breathing Microwave Plasma CAThode (AMPCAT) is proposed to overcome this issue through the use of a microwave plasma discharge, producing an extracted current in the order of 1 A with 0.1 mg s-1 of air. In this paper, the effect of varying magnetic-field strength and topology is investigated by using an electromagnet coil, which reveals a significantly different behaviour for air compared to xenon. The extracted current with xenon increases by 3.9 times from the zero-field value up to a peak around 150 mT magnetic-field strength at the antenna, whereas an applied field does not increase the extracted current with air at nominal conditions. A non-zero magnetic-field with air is however beneficial for current extraction at reduced neutral densities. A distinct increase in extracted current is identified at low bias voltages with air for a field strength of around 50 mT at the internal microwave antenna, consistent across varying field topologies. The effect of a lowered magnetic-field strength in the orifice region is investigated through the use of a secondary coil, resulting in an extracted current increase of 25 % for a relaxation from 6 mT to 1 mT, and demonstrating the beneficial impact of a locally reduced field strength on electron extraction.

Air-breathing electric propulsion has the potential to enable space missions at very low altitudes. This study introduces to a 0D hybrid formulation for describing the coupled intake and thruster physics of an air-breathing electric propulsion prototype. Model derivation is then used to formally derive main system's key performance indicators and estimate the figure of merit for the design of rarefied flow air intakes. Achievable performance by conical intake shapes are defined and evaluated by Monte Carlo simulations. Influence of inlet flow variation is assessed by dedicated sensitivity analyses. The set of requirements and optimality conditions derived for the downstream plasma thruster suggest concept feasibility within an achievable performance range.

We present the AQUAJET propulsion system, a cathodeless, ambipolar thruster test bed operating on multiple propellants including water. It is based on Electron Cyclotron Resonance (ECR) at 2.45 GHz using a simple permanent magnet configuration of the plasma source. We discuss the theoretical background of the technology, our flexible modular design that allows testing of many thruster geometry configurations, and modelling work done in preparation for testing.

Air-breathing electric propulsion (ABEP) refers to a spacecraft in very-low Earth orbit (VLEO) harnessing upper atmospheric air as propellant for an electric thruster. This allows the orbital altitude to be maintained via drag-compensation, removing the need for on-board propellant storage and allowing a mission lifetime which is not limited by propellant capacity. A cathode (or neutraliser) is required for the high-specific impulse electrostatic thruster designs proposed for an ABEP application. One such study is the AETHER EU H2020 project, which aims to design an ABEP system that can be tested on-ground in a VLEO-representative environment. There is therefore a need to develop a cathode for ABEP as conventional thermionic hollow cathodes are susceptible to oxygen poisoning. The Air-breathing Microwave Plasma CAThode (AMPCAT) presented here is based on a plasma electron source, using a 2.45 GHz microwave antenna directly-inserted into the plasma volume to ionise neutral air particles. This study details the cathode design and the results of iterative standalone testing, with a particular focus on: (a) the identification of a dual-mode current emission, with transition from lower-to higher-current mode with air at bias values around 70 V between the extracting anode and internal cathode surfaces, (b) a comparison of performance relative to xenon, for which the peak extracted current is 30–40% higher than air at equivalent inputs, and (c) the effect of antenna electrical isolation, using alumina shielding thicknesses in the 0.1–0.7 mm range. Standalone cathode tests demonstrate 0.8 A of stable extracted current with 0.1 mg/s mass flow rate of a 0.48O2 + 0.52N2 mixture, relative bias of 80 V and input microwave power of 70 W. To the authors' knowledge, the demonstration of an extracted current in the 1 A order using air, without visible material degradation after several hours of operation, is a novel development in the cathode literature.

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.

Air-breathing electric propulsion (ABEP) enables long duration missions at very low orbital altitudes through the use of drag compensation. A system-level spacecraft model is developed, using the interaction between thruster, intake and solar arrays, and coupled to a calculation of the drag. A quadratic solution is found for specific impulse and evaluated to identify the thruster performance required for drag-compensation at varying altitudes. An upper altitude limit around 190 km is based on a minimum thruster propellant density, resulting in required thruster performance values of 𝐼𝑠𝑝 > 3000 s and 𝑇 ∕ 𝑃 > 8 mN/kW for a realistic ABEP spacecraft. The orbit of an air-breathing spacecraft is propagated with time, which highlights the prescribed orbit eccentricity due to non-spherical gravity and therefore an increased variability in the atmospheric conditions. A thruster control law is introduced which avoids a divergent altitude behaviour by preventing thruster firings around the orbit periapsis, as well as adding robustness against atmospheric changes due to season and solar activity. Through the use of an initial frozen orbit, thruster control and an augmented 𝑇 ∕ 𝑃 , a stable long-term profile is demonstrated based on the performance data of a gridded-ion thruster tested with atmospheric propellants. An initial mean semi-major axis altitude of 200 km relative to the equatorial Earth radius, a spacecraft mass of 200 kg, 𝐼𝑠𝑝 = 5455 s and 𝑇 ∕ 𝑃 = 23 mN/kW, results in an altitude range of around 10 km at altitudes of 160–183 km during a period of medium to high solar activity.

A zero-dimension plasma model of thermionic hollow cathodes is here derived. The equations of conservation of mass, current and power are solved within the insert and orifice regions. The model introduces some elements of novelty compared to prior literature such as the inclusion of the electron current collected on the orifice lateral walls, and ion and electron currents collected on the orifice plate walls in the overall balance equations. Estimation of the power deposited on walls due to electron and ion bombardment in both insert and orifice regions are performed, thus, enabling comparisons among different geometrical configurations and operating regimes. The model results are compared against published experimental data and a thorough investigation of the model response to variation of geometrical and operational condition of a sample cathode are presented. A power budget, which includes power consumption and power deposition, of a sample thermionic cathode is also discussed.

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.

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.

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.

The work presented in this paper addresses specific issues of plasma cathodes operated on alternative propellants aiming at improving the current extraction of the technology using a novel DC plasma neutraliser developed by a partnership between the Surrey Space Centre and the Institute for Plasma Science and Technology of the Italian Research Council. This neutraliser uses thin diamond films on metallic disks (the cathode). The experimental development of the diamond-based plasma neutraliser (DBC) is described along with the preliminary characterisation and the resulting operating envelope. The highest achieved performance by the DBC was 0.215 A at 125 W and 10 sccm (0.67 mg/s) mass flow rate of krypton. Efficiency metrics (such as electron extraction power efficiency, and gas utilisation factor) for the neutraliser were compared for different diamond cathodes with unique film properties and operating conditions.

The Halo thruster is a low-power plasma propulsion concept, currently under investigation and development within the Surrey Space Centre at the University of Surrey in collaboration with Surrey Satellite Technology Ltd, Airbus DS and Imperial College London. The device is based on the electrostatic acceleration of propellant ions produced in a DC-powered magnetized plasma discharge characterized by a closed-loop electron drift sustained by the combination of electric and magnetic fields. Current research and development activities include: (i) experimental testing of different laboratory models to optimize the thruster performance in the 100 – 200 W power range; (ii) detailed plasma measurements to determine the underlying plasma physics; (iii) implementation of a plasma model for hollow cathode design; (iv) design and manufacturing of an optimized Halo thruster Engineering Model, including a tailored hollow cathode. This paper presents an overview of the aforementioned activities.

The Air-breathing Microwave Plasma CAThode (AMPCAT) has been developed for air-breathing electric propulsion in very-low Earth orbit. In this study, the standalone AMPCAT plasma characteristics are analyzed by means of several diagnostic tools and operation on xenon is compared to a conventional hollow cathode. A transition of AMPCAT extracted current from a lower (⁠ < 0.1 A) to higher-current (⁠ > 0.5 A) mode, triggered by increasing the negative cathode bias voltage, is accompanied by a significant rise in internal electron density and external electron temperature. The AMPCAT is coupled with a cylindrical Hall thruster in the 100–300 W power-level running on 0.5–0.7 mg/s of xenon, and the thrust is directly measured for cathode operation with both xenon and air. Stable thruster operation is demonstrated for the AMPCAT running on both propellants. For xenon, the performance is compared to a hollow cathode, which reveals matching discharge current profiles but a significantly higher thrust for the AMPCAT at low discharge voltages, approximately two times higher at 200 V. Langmuir probe measurements highlight a 30–40 V lower plasma potential in the cathode vicinity for the AMPCAT with xenon compared to both the hollow cathode and AMPCAT with air. This indicates a significantly improved coupling of cathode electrons to the thruster discharge, yielding an increased degree of ionization. Faraday probe and Wien filter results show that a larger current utilization efficiency drives the observed performance difference at low discharge voltages, rather than a significant change in ion acceleration or plume divergence.

The ion plume of a 72mm diameter Hall Effect Thruster operated on mixtures of xenon/nitrogen and xenon/air is investigated by means of a Wien filter (or E x B probe). The dependence of the velocities of the plume ions (Xe⁺, Xe²⁺, Xe³⁺, O₂⁺, O⁺, N₂⁺ and N⁺) on the operating parameters of the thruster (anode voltage, anode power, mass flow rate and magnetic field) is explored. The most probable ion acceleration voltages, the ion current and density fractions of the multi-propellant, multi-species ion beam, are computed from the Wien filter spectra through a dedicated post-processing analysis. The knowledge of these properties is fundamental for understanding the contribution of each ion species to the propulsive performance metrics of the thruster when operated on these molecular gas mixtures.

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 present the development of a steady state plasma flow reactor to investigate gas phase physical and chemical processes that occur at high temperature (1000 < T < 5000 K) and atmospheric pressure. The reactor consists of a glass tube that is attached to an inductively coupled argon plasma generator via an adaptor (ring flow injector). We have modeled the system using computational fluid dynamics simulations that are bounded by measured temperatures. In situ line-of-sight optical emission and absorption spectroscopy have been used to determine the structures and concentrations of molecules formed during rapid cooling of reactants after they pass through the plasma. Emission spectroscopy also enables us to determine the temperatures at which these dynamic processes occur. A sample collection probe inserted from the open end of the reactor is used to collect condensed materials and analyze them ex situ using electron microscopy. The preliminary results of two separate investigations involving the condensation of metal oxides and chemical kinetics of high-temperature gas reactions are discussed.

Several techniques have been developed recently for performing time-resolved laser-induced fluorescence (LIF) measurements in oscillating plasmas. One of the primary applications is characterizing plasma fluctuations in devices like Hall thrusters used for space propulsion. Optical measurements such as LIF are nonintrusive and can resolve properties like ion velocity distribution functions with high resolution in velocity and physical space. The goals of this paper are twofold. First, the various methods proposed by the community for introducing time resolution into the standard LIF measurement of electric propulsion devices are reviewed and compared in detail. Second, one of the methods, the sample-hold technique, is enhanced by parallelizing the measurement hardware into several signal processing channels that vastly increases the data acquisition rate. The new system is applied to study the dynamics of ionization and ion acceleration in a commercial BHT-600 Hall thruster undergoing unforced breathing mode oscillations in the 44–49 kHz range. A very detailed experimental picture of the common breathing mode ionization instability emerges, in close agreement with established theory and numerical simulations.

Human spaceflight to/on/from the Moon will benefit from exploitation of various in-situ resources such as water volatile and mineral. Evidence for water ice in Permanently Shadowed Regions (PSRs) on the Moon is both direct and indirect, and derives from multiple past missions including Lunar Prospector, Chandrayaan-1 and LCROSS. Recent lunar CubeSats missions proposed through the Space Launch Systems (SLS) such as Lunar Flashlight, LunaH-Map and Lunar Ice-Cube, will help improve our understanding of the spatial distribution of water ice in those lunar cold traps. However, the spatial resolution of the observations from these SLS missions is on the order of one to many kilometres. In other words, they can miss smaller (sub-km) surficial deposits or near-surface deposits of water ice. Given that future lunar landers or rovers destined for PSRs will likely have limited mobility (but improved landing precision), there is a need to improve the spatial accuracy of maps of water ice in PSRs. The VMMO (Volatiles and Mineralogy Mapping Orbiter) is a semiautonomous, low-cost 12U lunar Cubesat being developed by a multi-national team funded through European Space Agency (ESA) for mapping lunar volatiles and mineralogy at relatively high spatial resolutions. It has a potential launch in 2023 as part of the ESA/SSTL lunar communications pathfinder orbiter mission. This paper presents the work carried out so far on VMMO concept design and development including objectives, profile, operations and spacecraft payload and bus.

Water ice and other volatile compounds found in permanently shadowed regions near the lunar poles have attracted the interests of space agencies and private companies due to their great potential for in-situ resource utilization and scientific breakthroughs. This paper presents the mission design and trade-off analyses of the Volatile Mineralogy Mapping Orbiter, a 12U CubeSat to be launched in 2023 with the goal of understanding the composition and distribution of water ice near the lunar South pole. Spacecraft configurations based on chemical and electric propulsion systems are investigated and compared for different candidate science orbits and rideshare opportunities.

AVS-led projects are advancing the development of non-invasive diagnostic technology for Electric Propulsion (EP) thrusters. Such technology will gain importance as EP becomes an increasingly prevalent form of spacecraft propulsion. It is expected that future development will enable marketable products by early next decade. As such, the Beam Induced Fluorescence method has been assessed for its applicability to non-invasive diagnostics for EP. The concept was proved feasible by the ‘BIFEP’ project (a joint collaboration with Surrey Space Centre and support of the UK Space Agency). A further development called ‘ORBITA’ in collaboration with ESA will provide valuable datasets of performance parameters during operation of high power EP systems in-orbit and on-board the spacecraft.

We report on the results of an experimental campaign to measure time-varying velocity distributions in the near-field of a low power Hall thruster. We employ a sample-hold technique, enhanced by parallelizing the measurement hardware into several signal processing channels that vastly increases the data acquisition rate. The measurements are applied to study flow field dynamics in a commercial BHT-600 Hall thruster undergoing unforced breathing mode oscillations in the 44–49 kHz range. A very detailed experimental picture of the near-field emerges from these studies. The results indicate that velocity fluctuations lessen further downstream of the exit plane. Along the thruster axis where there is a general appearance of a central jet, there is evidence of a low velocity ion population in between the periodic bursts of high velocity ions, indicative of local ionization of neutrals outside of the thruster. One possible source of this residual ionization may be background chamber gas, which is not unexpected with the limited pumping capacity of ground test facilities.