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Pulsed plasma thrusters (PPT) have demonstrated enormous potential since the 1960s. One major shortcoming is their low thrust efficiency, typically

The inductively-heated plasma generator IPG6-B at Baylor University has been established and characterized in previous years for use as a flexible experimental research facility across multiple applications. The system uses a similar plasma generator design to its twin-facilities at the University of Stuttgart (IPG6-S) and the University of Kentucky (IPG6-UKY). The similarity between these three devices offers the advantage to reproduce results and provides comparability to achieve cross-referencing and verification. Sub- and supersonic flow conditions for Mach numbers between M a = 0 . 3 − 1 . 4 have been characterized for air, argon, helium and nitrogen using a pitot probe. Overall power coupling efficiency as well as specific bulk enthalpy of the flow have been determined by calorimeter measurements to be between η = 0 . 05 − 0 . 45 and h s = 5 − 35 MJ kg−1 respectively depending on gas type and pressure. Electron temperatures of T e = 1 − 2 eV and densities n e = 1 0 1 8 − 1 0 2 0 m−3 have been measured using an electrostatic probe system. At Baylor University, laboratory experiments in the areas of astrophysics, geophysics as well as fundamental research on complex (dusty) plasmas are planned. The study of fundamental processes in low-temperature plasmas connects directly to electric propulsion systems. Most recent experiments include the study of dusty plasmas and astrophysical phenomena and the interaction of charged dust with electric and magnetic fields. In this case, dust can be used as a diagnostic for such fields and can reveal essential information of the magneto-hydrodynamics in low-temperature plasmas. Although some of these goals require further advancement of the facility, it is proposed that several phenomena relevant to electric propulsion as well as to other fields of plasma physics can be studied using the existing facility.

In this paper, a generic full-system estimation software tool is introduced and applied to a data set of actual flight missions to derive a heuristic for system composition for mass and power ratios of considered sub-systems. The capability of evolutionary algorithms to analyse and effectively design spacecraft (sub-)systems is shown. After deriving top-level estimates for each spacecraft sub-system based on heuristic heritage data, a detailed component-based system analysis follows. Various degrees of freedom exist for a hardware-based sub-system design; these are to be resolved via an evolutionary algorithm to determine an optimal system configuration. A propulsion system implementation for a small satellite test case will serve as a reference example of the implemented algorithm application. The propulsion system includes thruster, power processing unit, tank, propellant and general power supply system masses and power consumptions. Relevant performance parameters such as desired thrust, effective exhaust velocity, utilised propellant, and the propulsion type are considered as degrees of freedom. An evolutionary algorithm is applied to the propulsion system scaling model to demonstrate that such evolutionary algorithms are capable of bypassing complex multidimensional design optimisation problems. An evolutionary algorithm is an algorithm that uses a heuristic to change input parameters and a defined selection criterion (e.g., mass fraction of the system) on an optimisation function to refine solutions successively. With sufficient generations and, thereby, iterations of design points, local optima are determined. Using mitigation methods and a sufficient number of seed points, a global optimal system configurations can be found., Ministerium für Wirtschaft, Arbeit und Wohnungsbau Baden-Württemberg, Projekt DEAL

This work investigates atomic recombination of gas species impinging on high-temperature materials used for thermal protection system of spacecraft entering Earth’s atmosphere. Heterogeneous cataly...

A common practice in the field of differential lift and drag controlled satellite formation flight is to analytically design maneuver trajectories using linearized relative motion models and the constant density assumption. However, the state-of-the-art algorithms inevitably fail if the initial condition of the final control phase exceeds an orbit and spacecraft-dependent range, the so-called feasibility range. This article presents enhanced maneuver algorithms for the third (and final) control phase which ensure the overall maneuver success independent of the initial conditions. Thereby, all maneuvers which have previously been categorized as infeasible due to algorithm limitations are rendered feasible. An individual algorithm is presented for both possible control options of the final phase, namely differential lift or drag. In addition, a methodology to precisely determine the feasibility range without the need of computational expensive Monte Carlo simulations is presented. This allows fast and precise assessments of possible influences of boundary conditions, such as the orbital inclination or the maneuver altitude, on the feasibility range., Projekt DEAL

A promising aspect of inductive plasma generators for space propulsion applications is their electrodeless nature that leads to a high propellant compatibility and opens the door to in-situ resource utilisation propellants. To further develop inductive plasma generators as a space propulsion technology it is important to understand the relationship between system inputs (e.g. input power and propellant material) and system outputs (e.g. power efficiencies) to aid design and control. In this work, methodologies are developed for non-intrusively assessing the inductive coupling from antenna current measurements in terms of the ‘instantaneous ignition power’, the ‘inductive duty cycle’ and the ‘inductive frequency shift’. Experimental results are presented for IPG7, a high-power (up to 180 kW) 5.5-turn helical antenna inductive plasma generator. Potential in-situ resource utilisation materials oxygen and carbon-dioxide, and their respective mixtures with argon, are assessed in terms of their inductive coupling characteristics and resulting power efficiencies. Oxygen is shown to be an attractive propellant material, especially when supplemented with argon. In particular, high thermal efficiency ( ∼ 84 % ) can be achieved at high input powers. A basis is provided for understanding the power efficiencies of an inductive plasma generator in terms of non-intrusive antenna current measurements, to aid power supply sizing and to develop monitoring and control techniques for thruster applications.

This study focused on the development of a vacuum cathode arc thruster (VAT), particularly on its design, manufacturing process, and demonstration. Characteristically, the proposed thruster does not require any additional propellant feeding system as the cathode electrode is simultaneously used as a propellant. In the ignition system, tiny spots are coated on the cathode surface to induce plasma flow. Such a setup has the advantages of simplicity, low price, small size, and low weight and is suitable for microsatellites. Moreover, a “trigger-less” method with an inductor storage power system was used for generating the pulsed plasma. This discharge method can significantly reduce input power. Thrust is mainly caused by high exhaust velocities of metal ions in the plasma flow, making ion density, ion velocity, and ion charge influential parameters. A battery was used instead of a power supply system to reduce the energy consumption of the entire VAT processing unit. The energy required for a single pulse was estimated to be 0.266 J, by measuring the change between the discharge current and the voltage. The ion current was measured using an ion detector and was 3.55 A, and the ion velocity was 23150 m/s. In the theoretical analysis, the VAT prototype proposed in this study achieved a single impulse of 4.31 μNs, a specific impulse of 1571 s, and a thrust efficiency of approximately 12.5%.

Several small, unconnected satellites flying in formation are a frequently discussed option to replace large satellites due to advantages such as increased flexibility, reduced costs and enhanced reliability. Using differential aerodynamic forces as a means to maintain the formation despite present perturbations, to perform formation reconfiguration or to initiate a rendezvous maneuver is a promising propellant-less alternative to chemical and/or electric propulsion systems. Subjected to uncertainties, dynamic variations and non-linearities its implementation in a real mission is, however, challenging. A common practice to either design suitable reference trajectories for robust control strategies or to gain deeper insights into the methodology is to use linearized relative motion models and a constant density assumption. Following this practice, multiple previous studies developed and enhanced analytic control algorithms to zero out the relative position and velocity of two spacecraft utilizing differential aerodynamic forces. Even though aerodynamic lift expands the controllability to all three translational degrees of freedom, it has been frequently neglected due to the low lift coefficients experienced in orbit so far. And even if considered, analysis has shown that the feasibility range of using the differential lift for the eccentricity control of the in-plane motion is extremely limited compared to using differential drag. However, two different possible options, namely applying enhanced algorithms to decrease the eccentricity before even initializing the eccentricity control phase as well as using advanced satellite surface materials targeting specular or quasi-specular reflection and thereby increasing the differential acceleration forces, have been suggested to increase the feasibility domain of the methodology. In this article, the effectiveness of the proposed adjustments is analyzed in a Monte-Carlo approach by applying them to the analytic control algorithms introduced in the literature. Since these are based on the linearized Schweighart-Sedwick equations, so is the presented analysis. Moreover, additional modifications to the algorithms are made to reduce the maneuver time of one of the respective control phases. The results show that the analyzed options noticeably improve the success rates of the maneuvers. Especially the feasibility range of the lift-based control algorithms is enlarged considerably. In addition, the implemented modifications are shown to consistently reduce the maneuver time of the respective control phase. Consequently, the presented analysis improves the current state of the art of the analytically designed rendezvous trajectories and provides valuable new insights into the methodology of using differential aerodynamic forces as a means of relative motion control.

This paper presents the results of an experimental campaign to measure thruster-relevant parameters for a high-power (180kW) inductive propulsion system utilising Ar, $ {\textrm{O}}_{2}$ , $ \textrm{N}_{2}$ , and $ \textrm{CO}_{2}$ as propellants. Results from the investigation show that inductive thrusters can make use of these propellants without the severe degradation seen in other electric propulsion systems. Furthermore, the collection of experimental data at powers greater than 100kW provides a reference of performance for the high-power electric propulsion devices intended for missions in the near future. Thrust and specific impulse in inductive systems can be improved by preferentially combining the chemical properties of atomic and molecular propellants. The maximum thrust recorded during these experiments was 7.9N, obtained using a combination of argon and oxygen (0.68 Ar + 0.32 $\textrm{O}_{2}$ ). The combination of argon and molecular propellants also decreased thermal losses within the discharge volume. Specific impulse can be doubled for the same input electric power by combining propellants, and future modifications to the thruster geometry and acceleration mechanism can be used to further improve the performance of such systems.

This article presents the full operational experimental capabilities of the plasma wind tunnel facilities at the Institute of Space Systems at the University of Stuttgart. The simulation of the aerothermodynamic environment experienced by vehicles entering the atmosphere of Earth is attempted using three different facilities. Utilizing the three different facilities, the recent improvements enable a unique range of flow conditions in relation to other known facilities. Recent performance optimisations are highlighted in this article. Based on the experimental conditions demonstrated a corresponding flight scenario is derived using a ground-to-flight extrapolation approach based on local mass-specific enthalpy, total pressure and boundary layer edge velocity gradient. This shows that the three facilities cover the challenging parts of the aerothermodynamics along the entry trajectory from Low Earth Orbit. Furthermore, the more challenging conditions arising during interplanetary return at altitudes above 70 km are as well covered., Projekt DEAL

The present study focuses on the influence of applied magnetic fields on argon plasma flows to assess the magnetohydrodynamics (MHD) effects on a pico-sized reentry capsule and their potential as a...

Characterization of IEC tight jet mode is achieved through Faraday probe measurements. Preliminary results indicate that the tight jet is a highly energetic electron beam with scattering of secondary ions and electrons, which result from electron beam impact ionization. A novel analytical model is proposed to evaluate this non-Maxwellian plasma beam, including a compensation of the secondary electron emission effect on the probe's surface. The results show prominent features on the extracted jet's kinetic energy and the respective (electron) current of IEC. In addition, they demonstrate a practical method to characterise non-Maxwellian plasma jets through Faraday probes.

Differential aerodynamic forces are a promising propellant-less option to control satellite formation flight. To this day, satellite lift is the most frequently neglected and, as a consequence, the methodology of differential lift only poorly studied. This is because the adsorption of atomic oxygen on the satellite’s surfaces in Very Low Earth Orbit induces diffuse reflection and high levels of energy accommodation, both of which results in the low lift coefficients experienced in-orbit so far. Analysis has shown that surface materials which promote specular or quasi-specular reflections are able to strongly increase the magnitude of the available differential lift forces. An influence of advanced surface materials on respective maneuver sequences, however, has not yet been analyzed at all. In addition, the robustness of the differential lift-based controller proposed up to now is questionable and not able to cope with the occurring uncertainties and dynamic variations. This paper aims to address these two research gaps. To do so, a robust control approach based on Lyapunov principles developed by Perez and Bevilacqua for the differential drag-based control of the in-plane relative motion is used in a subsequent second control phase to control the out-of-plane relative motion via differential lift. In a successive second step, the influence of different levels of energy accommodation on the full rendezvous maneuver sequence is analyzed. The results show that even a modest reduction in energy accommodation strongly reduces the maneuver times as well as the resulting orbital decay. In all analyzed cases, the proposed control approach led to a successful rendezvous.

The objective of this work is to study the behavior of surface layers of high-temperature metals in their interaction with a powerful pulsed plasma flow produced by a high-power ablative pulsed plasma thruster. This plasma generator produces plasma flows with a directed velocity of (7–9) × 106 cm/s, an initial diameter of 1.5–2 cm, and a maximum number density of about 1018 cm–3, as well as a maximum power of 5 GW. The main measured values are the residual temperature of the tungsten specimens and the evaporated mass. Also, metallographic analysis of the specimens was performed. The basis of the research method is to analyze the experimental data with the help of a numerical model describing the heating and evaporation of the material upon absorption of pulsed energy fluxes taking into account the evaporation kinetics based on the Hertz–Knudsen expression. Based on the developed numerical model and the obtained experimental data, the kinetics of evaporation of tungsten at high power fluxes to the surface (up to 1 GW/cm2) is investigated.

ZURAM is an ablative carbon–phenolic thermal protection material, developed by the DLR, German Aerospace Center (DLR) in cooperation with the Institute of Space Systems, University of Stuttgart. In...

The present work investigates the formation and stability of different oxides on PM1000, an Oxide dispersion-strengthened (ODS) Ni-Cr superalloy, and on SiC coated C/C-SiC after being exposed to highly dissociated nitrogen and oxygen flows in a Plasma Wind Tunnel (PWT). PM1000 and C/C-SiC are candidate materials for re-entry vehicles Thermal Protection System (TPS) and were used e.g. on the European eXPErimental Re-entry Testbed (EXPERT). The surface characterisation covers the identification of the oxide layers formed on the specimen's surface via X-ray Diffraction (XRD) and Energy-Dispersive X-ray spectroscopy (EDX). Grain sizes are assessed from XRD patterns and different topographical features are identified for Cr2O3 and NiO on PM1000 samples from Secondary Electron Microscopy (SEM). Surface spectral and total normal emissivities of the different oxides are assessed showing a significant dependency on the oxide, which is verified with previous investigations on ground and with flight data from the FOTON capsule.

The inductively-heated plasma generator IPG6-B at Baylor University has been established and characterized in previous years for use as a flexible experimental research facility across multiple applications. The system uses a similar plasma generator design to its twin-facilities at the University of Stuttgart (IPG6-S) and the University of Kentucky (IPG6-UKY). The similarity between these three devices offers the advantage to reproduce results and provides comparability to achieve cross-referencing and verification. Sub- and supersonic flow conditions for Mach numbers between $Ma = 0.3 - 1.4$ have been characterized for air, argon, helium and nitrogen using a pitot probe. Overall power coupling efficiency as well as specific bulk enthalpy of the flow have been determined by calorimeter measurements to be between $\eta = 0.05 - 0.45$ and $h_s = 5- 35$ MJ/kg respectively depending on gas type and pressure. Electron temperatures of $T_e = 1 - 2$ eV and densities $n_e = 10^{18} - 10^{20} m^{-3}$ have been measured using an electrostatic probe system. At Baylor University, laboratory experiments in the areas of astrophysics, geophysics as well as fundamental research on complex (dusty) plasmas are planned. Most recent experiments include the study of dusty plasmas and astrophysical phenomena and the interaction of charged dust with electric and magnetic fields. In this case, dust can be used as a diagnostic for such fields and can reveal essential information of the magneto-hydrodynamics in low-temperature plasmas., Comment: Accepted for publication in Acta Astronautica

Challenging space mission scenarios include those in low altitude orbits, where the atmosphere creates significant drag to the S/C and forces their orbit to an early decay. For drag compensation, propulsion systems are needed, requiring propellant to be carried on-board. An atmosphere-breathing electric propulsion system (ABEP) ingests the residual atmosphere particles through an intake and uses them as propellant for an electric thruster. Theoretically applicable to any planet with atmosphere, the system might allow to orbit for unlimited time without carrying propellant. A new range of altitudes for continuous operation would become accessible, enabling new scientific missions while reducing costs. Preliminary studies have shown that the collectible propellant flow for an ion thruster (in LEO) might not be enough, and that electrode erosion due to aggressive gases, such as atomic oxygen, will limit the thruster lifetime. In this paper an inductive plasma thruster (IPT) is considered for the ABEP system. The starting point is a small scale inductively heated plasma generator IPG6-S. These devices are electrodeless and have already shown high electric-to-thermal coupling efficiencies using O 2 and CO 2 . The system analysis is integrated with IPG6-S tests to assess mean mass-specific energies of the plasma plume and estimate exhaust velocities.

An experimental methodology to quantify atomic recombination on high-temperature materials exposed to homonuclear diatomic gases is proposed. Candidate materials are tested in a plasma wind tunnel ...

Electric propulsion offers the advantage of a high specific impulse through a large exhaust velocity and has seen significant progress in space flight applications. Recently, we observed a transient plasma shockwave during pulsed plasma thruster operation when the plasma beam impacted a probe surface. However, details regarding the plasma shockwave formation are still unknown. This work is an experimental investigation of the compression-induced plasma shockwave in the presence of a planar obstruction. To study the complete shockwave buildup and dissipation process, an ultra-high-speed imaging system was set up to visualize the time-resolved shockwave morphology at a sub-microsecond level. In addition, the local magnetic field and plasma density were measured using 2D magnetic coils and a triple Langmuir probe, respectively. The successive images of the shockwave give us a comprehensive understanding of the shockwave buildup process. During the 12 μs operational period of the thruster, two shockwaves were formed during the first cycle of the discharge. It is also interesting to note that there is a 1 μs dissipation period between the two shockwaves with the same cloud of plasma compressing against the probe surface. A shockwave model is also developed to predict the appearance of the two shockwaves. The implication is that the local magnetic field strength can be a key indicator for the plasma shockwave buildup and dissipation process.

The inertial electrostatic confinement (IEC) has been commonly researched in the scope of application for space propulsion and fusions reactors since many years. The utilization as thruster is possible due to a free emitting stream in a so-called jet mode of the IEC source. Depending on the setup, this jet can be operated as high energetic electron beam (tight jet) with electron energies of several keV or as low energetic plasma jet close to quasineutrality (spray jet). These modes are of high interest for thin film applications and plasma treatment. However, the IEC source is still not used in this field of application. Relevant application scopes can be electron substrate heating, surface plasma pretreatment, ionized physical vapor deposition and plasma enhanced chemical vapor deposition. In this paper, a cylindrical IEC source is presented that emits the plasma in both known jet modes, spray jet and tight jet. The jet emission occurs in radial outlet direction along the entire source height (“curtain-like” emission). Furthermore, an ignition test and operational behavior test is introduced. A stable spray jet operation region was examined and dependencies on pressure and voltage are investigated.

The results of a lunar massdriver mission and system analysis are discussed and show a strong case for a permanent lunar settlement with a site near the lunar equator. A modular massdriver concept is introduced, which uses multiple acceleration modules to be able to launch large masses into a trajectory that is able to reach Earth. An orbital mechanics analysis concludes that the launch site will be in the Oceanus Procellarum a flat, Titanium rich lunar mare area. It is further shown that the bulk of massdriver components can be manufactured by collecting lunar minerals, which are broken down into its constituting elements. The mass to orbit transfer rates of massdriver case study are significant and can vary between 1.8 kt and 3.3 megatons per year depending on the available power. Thus a lunar massdriver would act as a catalyst for any space based activities and a game changer for the scale of feasible space projects.

The Inertial Electrostatic Confinement (IEC) method is a plasma confinement principle with different operation modes. The work presented investigates the jet mode, which occurs with a jet extraction relevant for space propulsion applications. Two different jet modes were observed, the tight jet mode and the spray jet mode. Depending on propellant, discharge pressure and power these modes occur at different operation conditions and characteristics. Emission spectroscopic investigations show a different plasma species composition of confinement and extraction plasma and the transition between those modes has been examined with a high-speed camera. Finally, the applicability of IEC plasma sources for space propulsion systems is discussed with respect to the experimental results.

Microgravity experiments are essential for research in space science, biology, fluid mechanics, combustion, and material sciences. One way to conduct microgravity experiments on Earth is by using drop tower facilities. These facilities combine a high quality of microgravity, adequate payload masses and have the advantage of virtually unlimited repeatability under same experimental conditions, at a low cost. In a collaboration between the Institute of Space Systems (IRS) at the University of Stuttgart and Baylor University (BU) in Waco, Texas, a new drop tower is currently under development at the Center for Astrophysics, Space Physics and Engineering Research (CASPER). The design parameters of the drop tower ask for at least 1.5 s in free fall duration while providing a quality of at least 10 −5 g. Previously, this quality has only been achieved in vacuum drop tower facilities where the capsule experiences virtually zero aerodynamic drag during its free fall. Since this design comes at high costs, a different drop tower design concept, which does not require an evacuated drop shaft, was chosen. It features a dual-capsule system in which the experiment capsule is shielded from aerodynamic forces by surrounding it with a drag shield during the drop. As no other dual-capsule drop tower has been able to achieve a quality as good as or better than 10 −5 g previous work optimized the design with an aerodynamic perspective by using computational fluid dynamics (CFD) simulations to determine the ideal shape and size of the outer capsule and to specify the aerodynamically crucial dimensions for the overall system. Experiments later demonstrated that the required quality of microgravity can be met with the proposed design. The main focus of this paper is the mechanical realization of the capsule as well as the development and layout of the surrounding components, such as the release mechanism, the deceleration device and the drop shaft. Because the drop tower facility is a complex system with many interdependencies between all of the components, several engineering challenges had to be addressed. For example, initial disturbances that are caused by the release mechanism are a common issue that arises at drop tower facilities. These vibrations may decrease the quality of microgravity during the initial segment of free fall. Because this would reduce the free fall time experiencing high quality microgravity, a mechanism has been developed to provide a soft release. Challenges and proposed solutions for all components are highlighted in this paper.

Silicon Carbide (SiC) is a low Z material discussed as an alternative to graphite for fusion devices. The retention of hydrogenous species is an important plasma-surface interaction property. Deuterium was implanted into SiC, SiCf/SiC, Cf/C-SiC and SiC coated graphite under various particle energy and substrate temperature conditions. A TDS process was used to characterise the deuterium retention of the implanted specimens. While all SiC materials show elevated retention levels compared to graphite, the differences are limited to about a factor of two over the range of parameters investigated. Keywords: Deuterium Retention, Silicon Carbide, Ceramic Composite Materials, Low Z Materials

In this paper, results of the simulation tools URANUS and PICLas are compared to Plasma wind tunnel experiments. This is done for validation purposes of catalytic modelling within these tools. Therefore, the used experimental setup, from which the discussed simulation case is derived, and challenges in the flow characterization are shown. The main focus is on the validation of the newly developed surface Monte Carlo reconstruction method implemented in PICLas. For this method a brief description is given. It shows good agreements to experiments without a priori fitting of simulation parameters to experiments. However, further validation for evaluation of this method is needed.

To investigate the dissociation and recombination processes in a plasma jet, the characterization of its composition is important. For this purpose, an automated fitting algorithm for emission spectroscopic measurements, based on the radiation simulation tool PARADE, is developed. The algorithm is demonstrated by qualitatively analyzing single carbon dioxide spectra of the inductive plasma generator IPG4, operated at 125 kW anode power with a CO2 mass flow rate of 2.2 g/s. The species O, C, CO, CO+, CN and C2 are fitted with regards to number density, as well as vibrational and rotational temperature. A radial analysis of species composition is performed at two axial positions in the plasma jet. The results show different regions inside the plasma, suggesting increasing CO dissociation towards the centerline, as well as recombination of atomic carbon with molecular nitrogen. Moreover, the species composition distribution is found to be an indicator for the flow conditions in underexpanded free turbulent plasma jets.

Over 2500 active satellites are in orbit as of October 2020, with an increase of ~1000 smallsats in the past two years. Since 2012, over 1700 smallsats have been launched into orbit. It is projected that by 2025, there will be 1000 smallsats launched per year. Currently, these satellites do not have sufficient delta v capabilities for missions beyond Earth orbit. They are confined to their pre-selected orbit and in most cases, they cannot avoid collisions. Propulsion systems on smallsats provide orbital manoeuvring, station keeping, collision avoidance and safer de-orbit strategies. In return, this enables longer duration, higher functionality missions beyond Earth orbit. This article has reviewed electrostatic, electrothermal and electromagnetic propulsion methods based on state of the art research and the current knowledge base. Performance metrics by which these space propulsion systems can be evaluated are presented. The article outlines some of the existing limitations and shortcomings of current electric propulsion thruster systems and technologies. Moreover, the discussion contributes to the discourse by identifying potential research avenues to improve and advance electric propulsion systems for smallsats. The article has placed emphasis on space propulsion systems that are electric and enable interplanetary missions, while alternative approaches to propulsion have also received attention in the text, including light sails and nuclear electric propulsion amongst others., Irish Research Council

The further development of inductive plasma generators (IPGs) for space propulsion applications has driven the need for the development of non-intrusive experimental techniques that can resolve transient discharge cycle behaviour and accommodate a variety of propellant gases and gas-mixtures. In this work, a new approach is presented that simultaneously measures the antenna current, radially-resolved axial magnetic field strength and the radially-resolved emission in the visible range, at a sampling frequency sufficient to capture the driving and discharge cycle frequencies. Experiments were conducted using IPG7: a helical antenna IPG, driven nominally at 586 kHz, with 300 Hz discharge cycles. Nitrogen, argon-nitrogen, and argon-oxygen propellants were tested at capacitive and inductive conditions with input powers ranging from 4 kW to 36 kW. The propellants could be sourced from the space environment under in-situ resource utilisation. Analysis of the mean discharge cycle has made it possible to identify different coupling modes occurring within the cycle. These findings underscore the need to model such plasmas at the timescale of individual discharge cycles and demonstrate the value of the measured data for developing such models.