ZrB2 is a leading candidate for TPS in reentry vehicles owing to its high melting point, thermal shock resistance, and low density. It improves oxidation resistance when composited with SiC. Since TPS is used in the airframes of reusable space vehicles, it must possess excellent mechanical properties. To achieve this, the manufacturing process was optimized. ZrB2-SiC and TiB2-SiC with different composition ratios were fabricated through SPS. The microstructure was most densified sintered at 1900 degrees C and for 10 min. A comparison of the microstructures of TiB2-SiC and ZrB2-SiC mixtures revealed that the TiB2 particles had agglomerated and grown into coarse grains, indicating that the sintering conditions significantly affect microstructure formation. The ZrB2-30 vol%SiC and TiB2-15 vol%SiC mixtures had the highest Vickers hardness, at 2345(HV5) and 3223(HV5), respectively. These samples were selected for subsequent oxidation testing. For TiB2-15SiC, the depth of the oxidation reaction layer increased with oxidation time at 1250 degrees C. For ZrB2-30SiC, the oxide layer thickened with increasing oxidation time, and passive oxidation only occurred at 1250 degrees C. At 1650 degrees C, a non-oxidized ZrB2 layer was observed beneath the oxide layer, whereas the TiB2-15SiC sample shattered. ZrB2-SiC was proved to a promising material for TPS owing to its superior mechanical properties and excellent oxidation resistance.

Atmospheric re-entry represents a crucial stage in crewed spaceflight and sample-return missions, making it essential to understand plasma behavior around the vehicle. This paper presents a numerical rebuilding approach for high-enthalpy plasma flows in plasma wind tunnel (PWT) experiments, employing the axis-symmetric thermo-chemical non-equilibrium Navier–Stokes code URANUS developed at the University of Stuttgart. Test conditions with high enthalpies of 60 MJ kg ^-1 and 80 MJ kg ^-1 were characterized, corresponding to highly elliptical and hyperbolic re-entry trajectories, respectively. The numerical simulation results demonstrate excellent agreement with experimental measurements for both conditions, highlighting the importance of ion recombination at the vehicle’s surface. The validated numerical methodology established in this study will be utilized to characterize recent experiments employing a probe equipped with solenoid magnets to investigate magnetohydrodynamic (MHD) effects during atmospheric re-entry.

Satellites in very low Earth orbit (VLEO) decay over time due to drag. However, payloads can be benefited, potentially improving performance and costs. To enable VLEOs, the Institute of Space Systems (IRS) is developing an atmosphere-breathing electric propulsion (ABEP) system. The current design is based on an electrodeless RF helicon-based plasma thruster (IPT) laboratory prototype. It features a solenoid that provides the external magnetic field for plasma confinement and acceleration. Based on this, a vacuum capable model is developed, using as alternative a magnet configuration. Different ring magnet configurations are investigated, optimizing the magnetic field topology. Performance and plasma properties are extracted by a theoretical model. A simulation process approach is presented. A sensitivity analysis is conducted, identifying factors affecting the magnetic field topology. A simple way is proposed to change the field strength, keeping field topology and geometry the same by adjusting the internal diameter of the magnets. As a result, changes of the field strength and the resulting performance parameters, required by future developments of the IPT, can be accounted for.

Space missions in very low Earth orbit (VLEO) offer advanced capabilities for Earth observation, telecommunications, and security but are challenged by continuous orbit decay. Atmosphere-Breathing Electric Propulsion (ABEP), which uses atmospheric particles as propellant, provides a potential solution. Within the H2020 DISCOVERER, ESA ram-CLEP and ATLAS projects, IRS is developing a specular intake and helicon plasma thruster to advance ABEP systems. This study uses Direct Simulation Monte Carlo (DSMC) simulations to evaluate specular intake geometries, analyzing particle flux, pressure, mass flow rate, and collection efficiency. The results indicate that reducing the focal length of the parabolic intake significantly improves efficiency and mass flow rate, while increasing the discharge channel diameter proves more effective for achieving these high values as well. As pressure follows an opposing trend to efficiency, and reaching the ignition pressure is crucial, the optimal configuration among the investigated for balancing efficiency and pressure is a discharge channel diameter of 25 mm with a focal length of 3 mm.

The space industry aims to promote economic growth, while ensuring a sustainable use of space. The sustainability objectives of space often refer to Active Debris Removal missions, which, despite being needed, are still following a linear economic model. Space circular economy permits giving value to space debris, considering it as a valuable resource. Pilot studies investigated the application of circular concepts to spacecraft solar panels. The results of a first study on the thermal degradation behavior of the solar cell encapsulant, WACKER® RTV-S 691, identified a degradation starting at 300 ^∘ C. It demonstrates great potential for further investigation of thermal delamination processes. This paper aims to propose an experimental approach to investigate the use of solar radiation to delaminate and to anneal spacecraft solar cells using a Sun simulator with a maximum power of 1500 W m^-2 . This initial study aims to start the discussion of the implementation of circular concepts within in-space assembly and manufacturing to facilitate their implementation.

We report a new experimental platform developed at the DIII-D National Fusion Facility to investigate carbon ablation and spallation under extreme heat fluxes relevant to fusion plasma-facing components and high-enthalpy atmospheric entry. Carbon samples were exposed to parallel heat fluxes of $30$--$40~\mathrm{MW\,m^{-2}}$ in the scrape-off layer using two complementary approaches: stationary carbon rods inserted near the divertor strike point and slow-launch carbon pellets injected vertically into the edge and core plasma. Pellets penetrating the core experienced heat fluxes approximately an order of magnitude higher. The conditions reproduce key aspects of the shock-layer environment encountered by the Galileo probe during entry into Jupiter's atmosphere. Fast visible imaging, divertor spectroscopy, infrared thermography, CO$_2$ interferometry, and post-exposure profilometry provided measurements of ablation rates, surface recession, and temperature evolution. Measured mass-loss rates of $(1$--$3)\times10^{-2}~\mathrm{g\,cm^{-2}\,s^{-1}}$ agree with semi-empirical aerospace ablation models, while wedge-shaped rods exhibited greater ablation than cylindrical and concave samples. UEDGE-DUSTT simulations incorporating parallel plasma flows, ${\bf j}\times{\bf B}$ forces, and ablation-cloud shielding reproduce the measured pellet trajectories and ablation timescales. These results establish tokamak plasma as a high-heat-flux environment for validating carbon ablation models and studying material response and impurity dynamics in reactor-relevant divertor plasmas.

Pulsed, gas-fed electric propulsion systems are a promising option for future small satellite missions but pose significant challenges for accurate thrust characterization due to parasitic forces introduced by propellant feed lines. This paper presents and experimentally demonstrates an indirect thrust measurement methodology for a pulsed quasi steady-state magnetoplasmadynamic thruster prototype of the Modular Pulsed Propulsion System (\name). The thruster plume is directed onto a graphite target mounted on a high-precision thrust balance, decoupling the balance mechanically from the gas supply system. Since the impulse transfer between the argon plume and the target is unknown, it is quantified using LAMMPS molecular dynamics simulations to obtain an impulse transfer correction factor. The impulse bit is measured using nanometer-resolution pendulum displacement measurements processed via sinusoidal curve fitting and wavelet filtering. A comprehensive uncertainty analysis based on Gaussian error propagation is performed, accounting for calibration, sensor resolution, geometry, timing, frequency determination, damping, and impulse transfer. The methodology is used to determine impulse bit, specific impulse, and thrust efficiency of the \name, demonstrating thrust levels significantly above the cold-gas contribution alone, identifying distinct discharge modes, and outlining the design and operational changes required to approach the targeted performance.

This study, conducted as part of the EU Horizon 2020 project on Magnetohydrodynamic (MHD) Enhanced Entry Systems for Space Transportation (MEESST), investigates the heat flux mitigation potential in high enthalpy air plasma using a high-temperature superconducting solenoid within an MHD plasma probe. The effects of varying magnetic flux density on stagnation heat flux and pressure are examined, with an emphasis on altered surface conductivity. The high enthalpy air flow is generated using a self-field magnetoplasmadynamic plasma generator. The experiments demonstrate a reduction in total heat flux on the MHD plasma probe of up to 83 %, while limited effects of surface conductivity on stagnation heat flux and shock standoff distance are observed and correlated to the Hall parameter. These findings highlight the potential of MHD shielding in improving thermal protection for spacecraft during atmospheric entry.

The SUPREME (Superconductor-Based Readiness Enhanced Magnetoplasmadynamic Electric Propulsion) thruster, where the scaling approach shown here is performed by Neutron Star Systems in collaboration with the Institute of Space Systems (IRS), University of Stuttgart, is an applied field magnetoplasmadynamic thruster designed for a nominal power level of 5 kW to have a wide range of commercial applications. Basis for the design is the SX3 thruster, developed at IRS, which is an engineering model operated between 2 and 100 kW with magnetic fields up to 400 mT. For flight applications, copper coils, as used for the SX3, are limited in achievable magnetic field strength and suffer from high thermal losses. High temperature superconducting coils provide a path to flight technology, minimizing these losses and providing high magnetic field strengths like the 1.2 T intended for SUPREME, which will also have radiation cooled electrodes. This work describes a design approach in the absence of a numerical simulation code for the plasma inside the thruster. The downscaling process described here relies on maintaining critical plasma parameters of the SX3 thruster to produce similar thruster performance with the smaller SUPREME. The downscaling to obtain a discharge unit geometry and the design process for the superconducting coil are described in detail, the SUPREME performance is estimated with a scaling code developed at IRS. Lastly, potential mission scenarios within the Earth’s gravitational field and interplanetary missions to Mars and Ceres are investigated to assess potential applications for SUPREME.

The Institute of Space Systems is currently developing a deorbit module based on thermal arcjet technology to allow fast orbit decay at end-of-life, with a focus on megaconstellation satellites. By employing additive manufacturing with tungsten, improved nozzle geometries can lead to a gain in overall performance. However, reproducibility is an ongoing concern for additively manufactured parts. Together with the Heinz Maier-Leibnitz Zentrum (MLZ) and the Budapest Neutron Centre of the Centre for Energy Research, a study was conducted scanning additively manufactured arcjet nozzles prior to and after standardized operation via neutron computed tomography. The results show a drop in performance over time, which can be related to changes in the constrictor geometry. Furthermore, cavities created during manufacturing can significantly influence operation.

An electrostatic probe measurement has been established to measure the electron temperature, ion temperature, and electron density within an expanding plasma jet. Due to the plasma being in a transitional hydrodynamic regime, neither collisionless nor collisional probe theories could be directly applied to the present case. Therefore, an inverse method was used in which the shape of the current-voltage curve of the probe was iteratively calculated for given plasma conditions and directly compared with the measured result. Using this method, the temperatures and densities could be estimated even though the established Langmuir probe theory was not applicable.

Recently, BOOST (Building blOcks for iOdine thruSTers) a research initiative funded by the European Union (GA-101135216), kickstarted. The primary goal of BOOST is to advance and endorse iodine electric propulsion technology, positioning it as a crucial element to enhance the SmallSat market. In the last decade, SmallSats (under 500 kg) have transformed the space market by offering much more affordable access to space than medium-to-large systems. This trend is anticipated to grow in the coming years. Solid iodine propellant has shown potential as a key technology for establishing Electric Propulsion (EP) in the SmallSat sector. However, the broad adoption of this technology is still limited by its low maturity. To lay the foundation for future commercialization of iodine-based EP, the BOOST project will employ a modular approach in order to tackle the development of fundamental building blocks for this type of propulsion.

The Small Satellite Student Society at the University of Stuttgart (KSat e.V.) has a strong history of ferrofluid-based research, demonstratedthrough ISS projects PAPELL and FARGO, as well as the REXUS sounding rocket project FerrAS. The latest endeavor, projectFINIX, is scheduled to launch on the REXUS 34 sounding rocket in Q1 2025. Building on these previously conducted missions, FINIXseeks to advance ferrofluid-based solutions with future applications in space systems. These novel developments hold the potential toreplace conventional mechanical components in space applications, which are subject to wear and tear and hence represent a limitingfactor for space missions. This paper details the experiments planned for the technology demonstrator mission FINIX, focusing onthe development of an electrical switch and a pump, and their connection to earlier developments. The design, testing, and validationprocesses of these ferrofluid-based experiments will be discussed, highlighting the scientific objectives and innovations of the project.

The improvement of the prediction, reconstruction, and evaluation of radiative markers observed remotely during destructive atmospheric re-entries of de-orbited spacecraft relies on the availability of relevant reference datasets generated through ground-based testing. To this end, five distinct spacecraft materials, including grade 5 titanium Ti6Al4V, type 316L stainless steel, aluminium alloy 7075, aluminium-lithium alloy 2099, and Carbon Fibre-Reinforced Polymer (CFRP) EX-1515/M55J, were subjected to supersonic high-enthalpy air flow conditions in the plasma wind tunnel facilities of the University of Stuttgart. For two of the metal alloys, additional tests were conducted following the prior application of an Aeroglaze Z306 coating. Temporally resolved radiation responses were measured downstream of the stagnation point within the boundary layer and wake flow area, cross-examined with thermographic and visual data and discussed qualitatively. The results of the experiments indicate that material emissions, specifically those from metallic alloys, may provide a means of identifying the composition and possibly the ongoing state of demise of the observed materials under certain conducive circumstances. However, energy densities within the boundary layer near the stagnation area of an exposed sample appear to be generally insufficient to elicit clearly observable emissions associated with the primary element of a given metallic alloy, which was rarely identified in the measurements. This prompts a recommendation for specialised experimental setups to substantially increase droplet-specific energy densities within the measurement volume of the spectrographs deployed in ground testing.

Challenging space missions at very low altitudes face significant atmospheric drag, requiring efficient propulsion methods such as Atmosphere-Breathing Electric Propulsion (ABEP) to extend mission lifetimes. ABEP captures atmospheric particles and uses them as propellant for an electric thruster, reducing dependence on limited on-board propellant. This could extend missions in Very Low Earth Orbit (VLEO) and on celestial bodies with an atmosphere, such as Mars. The Institute of Space Systems (IRS), under the EU H2020 DISCOVERER, ESA Ram-CLEP, and CRC ATLAS projects, is developing a high-efficiency specular intake and a RF Helicon-based plasma thruster (IPT) for ABEP. This study uses the numerical tool PICLas and it’s Direct Simulation Monte Carlo Method’s (DSMC) to analyse the effect of solar activity and evaluate the validity of the hyperthermal assumption in VLEO for ABEP intake designs. Additionally, the effect of changing intake lengths on important key parameters, such as intake efficiency, mass flow rate, and pressure, is examined. The results show that efficiency decreases with higher solar activity, longer intakes and higher altitudes, with particle temperature having the greatest effect on efficiency, due to its influence on thermal velocity and the molecular speed ratio. An almost linear relationship between efficiency and molecular speed ratio is shown, revealing that the hyperthermal assumption may not be valid for VLEO applications. To achieve the required pressure level for ignition, flexible ABEP operation is recommended to accommodate for varying solar activity, suggesting lower altitude operation during low solar activity and higher altitude operation during high solar activity.

Applied-field magnetoplasmadynamic thruster (AF-MPDT) is a type of spacecraft electric propulsion (EP) with input powers of up to several 100 kW, which uses an external magnetic field to accelerate a propellant and to generate thrust. They reach a high-thrust density for an EP and medium to high specific impulses (Isp), which enables their usage in interplanetary space missions and for space tugs, for example, to the Moon. At the Institute of Space Systems, the water-cooled 100 kW steady state AF-MPDT SX3 is investigated. Magnetic field measurements with a Hall-probe were performed for a wide variation of thruster operating conditions, in order to determine the local current densities and, hence, the current densities’ distribution in the plume of the AF-MPDT SX3 for thruster distances between 140 and 1000 mm. Up to 70%–80% of the discharge current was measured as current in the plume at a thruster distance of 1000 mm. At a thruster distance of 300 mm, the fraction was 90% and more for the same operation conditions. An azimuthal current, with a current density of 7.9±2.2mA/mm2, was measured with an applied field of 100 mT at a thruster distance of 403 mm. Increasing applied magnetic fields moved it further downstream. Lorentz forces, calculated from the currents in the plume, are of the order of 10% of the thrust, which was measured with a thrust balance before.