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Abstract: ESA SysML Solution is a pack containing an MBSE methodology based on ECSS, a data model representation of the ECSS elements on SysML v1, an implementation in Cameo Systems Modeler and Enterprise Architect. In this paper, we present the evolution of the ESA SysML Solution presented in MBSE 2022 conference, focusing on verification and interface management. We address here Advancements, Limitations, and Customization of MBSE Approaches in Space. PubDate: 2025-06-30
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Abstract: A new approach for estimating the system response time (SRT) is presented and demonstrated through the analysis of the effect of target latitude on SRT, both with and without inter-satellite links (ISL). The access time intervals for each satellite to ground stations, targets, and between pairs of satellites are computed. The SRT for a target is calculated using a proposed algorithm that includes an adapted Dijkstra algorithm on a time-varying link graph for the ISLs. For the geospatial analysis, the average, minimum, and maximum SRTs of the targets of all SRTs were introduced as performance metrics. Combined with a visualization on a world map, this method allows us to quickly grasp the strengths and weaknesses of the constellation with regards to SRT. When plotted as a function of latitude [− 90°, 90°], the SRT for constellations with and without ISL in mid and high-inclination orbits peaks at the equator and drops towards the Earth’s poles, influenced by the satellite constellation and ground station network design. ISLs smoothen the SRT curve and reduce the effect of the latitude, thereby enhancing data acquisition speed by minimizing SRT variation over time and across longitudes. PubDate: 2025-06-30
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Abstract: Understanding the gas-surface momentum transfer of oxygen atoms, a major component of the residual atmosphere at very low Earth orbit (VLEO) altitudes of 100–450 km, on materials on the exterior of satellites is important for estimating drag. We have thus investigated the scattering dynamics of orbital-velocity O atoms on four representative materials fluorinated ethylene propylene (FEP) polymer, aluminum with a chromate conversion Alodine coating (Al), solar cell cover glass with a MgF2 coating (CG), and a glass-reinforced epoxy laminate circuit board material (FR4). A pulsed hyperthermal atomic-oxygen beam with a nominal translational energy of ~ 4.7 eV was directed at the target surface, and the scattered products were detected with a rotatable mass spectrometer. Time-of-flight (TOF) distributions were measured with various incident beam angles (θi = 60°, 45°, 30°, 15°, 0°) for O atoms scattered in and out of the plane defined by the incident beam and surface normal. For both in-plane and out-of-plane scattering experiments, TOF distributions of O atoms exhibited mostly impulsive scattering, with atom-surface scattering events occurring on a time scale too short for thermal equilibrium to be achieved. A small fraction of thermally desorbed O atoms was also observed. In addition, both the flux and energy of the scattered O atoms were found to be higher when exiting the CG and FEP surfaces than when exiting the FR4 and Al surfaces. The lower flux of O atoms scattering from FR4 and Al at a given final angle (θf), is the result of the combined effects of reactive collisions leading to OH and H2O products and the multiple-bounce trajectories of the incident O atoms on the rough surfaces, which leads to scattering angle randomization. Characterization of the material surfaces was consistent with the observed scattering dynamics: CG and FEP surfaces are relatively smooth at the nanometer scale, while FR4 and Al surfaces are comparatively rough. Regardless of surface roughness, the average fractional energy transfer to the surface is well parameterized by the angle through which the incident O atoms were deflected as they scattered from the surface. The fraction of thermally desorbed O atoms tends to be higher for the FR4 and Al surfaces than for the CG and FEP surfaces, which is consistent with greater energy accommodation on the FR4 and Al surfaces. The results suggest that FR4 and Al surfaces will lead to increased drag compared to CG and FEP surfaces, as a result of the greater overall energy accommodation of incident O atoms on the rougher surfaces. PubDate: 2025-06-27
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Abstract: Nuclear thermal propulsion (NTP) and nuclear electric propulsion (NEP) systems are considered to be potential enablers for exploring Mars and other outer planets. The fission surface power can provide continuous heat and electricity, which complements solar power on the Moon and Mars. Nuclear power sources for propulsion and electricity generation are attracting spacefaring nations and stakeholders again after the Space Race era due to their advantages and envisioned applications. This paper examines space nuclear power from sustainability perspectives while complying with safety requirements. First, from the space logistics and transportation perspectives, we identify potential operational regimes where nuclear space propulsion (NTP/NEP) could complement or outperform chemical rocket propulsion. Second, we discuss the lifecycle-wide design considerations of surface nuclear power as extraterrestrial infrastructures, inspired by the latest design approaches and technologies adopted by the architecture, engineering, and construction (AEC) industry as well as small modular reactor (SMR) manufacturers. The findings show that (i) nuclear-powered propulsion reduces the fuel consumption of space transportation in some cislunar missions and (ii) local sourcing of materials on the Moon could reduce the construction costs of surface nuclear facilities. Further assessment of safety and reliability will be a critical enabler to incorporate nuclear technologies into cislunar economy and beyond. PubDate: 2025-06-27
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Abstract: The simplicity, efficiency, and reliability of pulsed plasma thrusters make them a preferred choice in low-power systems for microsatellites. In this paper, a coaxial electrothermal pulsed plasma thruster was designed and tested using Teflon as a solid propellant. The ablated mass was measured for different values of peak discharge current and number of pulses. The Experimental results of our system were compared with calculated values obtained using a 1-D time-dependent ETFLOW model for the electrothermal capillary system. The model was used to calculate the values of thrust and impulse for each case. Experimentally, increasing the number of pulses resulted in an increase in the total ablated mass and the capillary diameter. Experimentally, increasing the peak discharge current from 1.95 kA to 5.66 kA led to an increase in the ablated mass per pulse from 128.4 µg to 980.8 µg and the impulse from 340 µN.s to 2156 µN.s. In addition, extending the capillary length from 3 to 8 cm led to an increase in the ablated mass from 52 µg to 326.2 µg, consistent with the calculated data. Moreover, the thrust efficiency was investigated as a function of input energy and capillary length. The maximum achieved thrust efficiency was about 8.9% for an input energy of 6 J and a capillary length of 6 cm. Theoretically, the model predicted that thrust, impulse, and specific impulse increase with raising the peak discharge current. PubDate: 2025-06-25
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Abstract: Over the last decade, the need for hypersonic systems has become a key priority for many government agencies across the world. However, flight-testing opportunities remain limited. Performance of hypersonic platforms is primarily examined through analytical models or ground testing, owing to the high cost and high risk of flight testing. The flight environment at hypersonic speeds is simulated with testing facilities such as wind tunnels, shock tubes, and plasma tunnels; however, limitations with each approach mean the full flight envelope is unable to be fully replicated on ground. Progress is slow as a result. With this renewed focus on hypersonic and suborbital platforms, new commercial flight-testing capabilities are entering the market such as Rocket Lab’s HASTE and DLR MORABA’s VS-50. The HiSST conference posed an opportunity to gather experts in the field to provide an overview of their platforms, to collect the needs and interests of the research and industrial community for their technologies’ demonstration, and to understand blockers for advancing particular concepts. PubDate: 2025-06-24
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Abstract: We describe the SMARTnet Instrument Enhancing Software (SMARTies) package, which we developed from scratch for fully automated remote control of our telescope stations within the Small Aperture Robotic Telescope Network (SMARTnet). Since March 2024, we have been using SMARTies continually to operate our SMARTnet telescope station in Chile successfully. We include a detailed description of the system design and architecture including the SMARTies modules, which are written in pure Python and are kept rather abstract, as well as the bespoke device controllers, which facilitate the communication with the actual hardware devices used. SMARTies was designed to fulfil a number of different use cases, including satellite or space debris tracking, survey observations, and observations of astronomical objects, including light curve acquisitions, and is therefore useful for the Space Situational Awareness community and other observational astronomers alike. While a fully automated observing mode following a user-defined schedule was one of the driving factors in the development of SMARTies, it does allow for near real-time manipulations of the schedule and even completely manual operations of the telescope. Because of its object-oriented and modular approach, new SMARTies functionalities can easily be added and different hardware devices can easily be included by adding new device controllers. Therefore, we envision SMARTies to be an extremely useful asset for many telescope operators across the world. PubDate: 2025-06-21
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Abstract: The growing deployment of nanosatellites in Low Earth Orbit has increased interest in aerodynamic control methods like differential drag. This technique adjusts a satellite’s attitude to vary drag and influence relative motion—without the need for propulsion. While theoretically flexible, real-world application is limited by operational constraints. Experiences from BEESAT-4 have shown that factors such as limited satellite availability and attitude control inaccuracies reduce the effectiveness of drag-based manoeuvres. This paper focuses on the S-NET formation—four 9 kg nanosatellites launched in 2018—and proposes a planning routine for practical differential drag control. The goal is to develop a flexible methodology that includes both operational constraints and uncertainties in environmental parameters. Potential manoeuvres are calculated using linearised relative motion equations and advanced models for aerodynamics and the orbital environment. By considering key limitations during planning—such as reduced availability and control accuracy—the impact of these constraints can be better understood. The analysis also accounts for uncertainties in aerodynamic parameters, offering a range of expected outcomes. The resulting planning tool supports the design of manoeuvre sequences that remain effective despite real-world challenges, offering valuable insights for future nanosatellite missions. PubDate: 2025-06-18
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Abstract: Space mission planning involves coupled architecture decisions that non-trivially influence system-level performance metrics such as weight, power usage, and scientific value. We present an exemplary space mission planning problem involving several mission-level and spacecraft-level choices, and optimizing for the conflicting objectives of system mass and scientific value. The problem is solved using System Architecture Optimization (SAO): a technique where numerical optimization algorithms are used to explore an architecture design space and find a Pareto front of optimal architectures. The architecture design space is modeled using the Architecture Design Space Graph (ADSG) implemented in the ADORE editing and optimization tool. The design space model includes function-component allocation choices, component-level design variables, and system-level objectives and constraints to optimize for. Evaluation code is implemented in Python and linked to the design space model using class factories. The design space is explored using NSGA-II, a multi-objective evolutionary algorithm, resulting in a Pareto front trading-off system mass and total experiment duration. PubDate: 2025-06-18
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Abstract: In VLEO, the continuum assumption breaks down requiring the use of the DSMC method to simulate the gas flow around satellites. However, DSMC simulations relies on simplified GSI models, such as the Maxwell model or the Cercignani–Lampis model, which are based on constant accommodation coefficients. Instead, these coefficients are variable, influenced by multiple factors, which makes their accurate determination challenging. Furthermore, these models are using simplifying assumptions, such as superposition of specular and diffuse reflections, or independent scattering of the normal and tangential components of the velocity. Implementing a high precision GSI model enables to optimize the aerodynamics of VLEO satellites and to design efficient intakes for atmospheric breathing propulsion systems vastly enhance mission planning and fuel requirement calculations, ultimately extending operational lifetimes and reducing costs. We present an approach that integrates MD simulation data into DSMC through a scattering kernel modeled as a Gaussian mixture conditional probability density function. Preliminary tests on synthetic data derived from the Cercignani–Lampis model demonstrate the model’s capability to accurately predict reflected velocities. Additional improvements are needed to enhance the model’s performance in interpolation and extrapolation of unknown incoming velocities in the future. PubDate: 2025-06-13
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Abstract: Lunar regolith is identified as one of the greatest challenges for future lunar surface missions. With a rising number of planned surface exploration missions within the next decades, the significance of technologically robust, dust-tolerant solutions and lunar dust mitigation strategies have never been more present. Within this context, the University of Stuttgart’s Institute of Space System is investigating technical solutions and methods for active and passive dust mitigation strategies. A low-fidelity test environment was set up in order to provide the ability to characterise adhesion forces and abrasive effects of lunar simulant particles on multiple technical surfaces. Microscopic imaging is used in order to quantify remaining particle depositions after several experiment series. The test campaigns presented in this paper include three specific experiments: dust particle adhesion characterisation on a specific surface type by centripetal force measurements (1), dust-surface cleaning and abrasion tests with varying types of brushes (2), as well as magnetic cleaning methods (3). A developed optical particle detection algorithm is being used in situ to the experiment to resource efficiently specify the amount and size of remaining particles to derive data on the respective adhesiveness. The low-fidelity test environment is seen as a very efficient solution for first precursor tests and experiment series with varying boundary conditions. The control and variation of sample surface substrates, lunar analogue types, grain sizes, and some environmental conditions allow a high variation of different testing scenarios in order to characterise first impact factor dependencies. This paper describes the general setup of the developed test environment, the specific experiments, methods and results, as well as the lessons learned. PubDate: 2025-06-11
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Abstract: The Dust Impact Sensor and Counter (DISC), part of the payload of Comet Interceptor mission, will determine the coma dust features of the mission target comet. DISC sensing plate will be exposed to cometary du... PubDate: 2025-06-04
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Abstract: The remarkable progress SpaceX made in the first four integrated flight tests of their Starship and Super Heavy launcher configuration indicates that a fully reusable space transport system might become a reality within a few years. Such a system could revolutionize the global launch market, especially if it is able to achieve its forecasted payload capacity of more than 100 t into LEO. Therefore, it is necessary to gain a deep understanding of the capabilities of this system and compare it to potential future European options for heavy and super-heavy space transport systems. This paper uses the publicly available data from Starship’s first four integrated flight tests for a thorough technical analysis of its current capabilities. The flight tests allow a calibration and update of our earlier-presented Starship models with real flight data. These updated models will be used to gain an understanding of its high-level system properties and to extrapolate the actual LEO capabilities of the early operational Starship versions. The second part of the paper will investigate a potential European option for launching similar payloads of 50 t and more into LEO, based on building blocks currently proposed or under investigation. This configuration employs a reusable winged first stage based on the SpaceLiner concept’s booster stage. The stage uses cryogenic liquid hydrogen and cryogenic liquid oxygen and is recovered with in-air capturing (IAC). For the second stage, an expendable cryogenic stage is optimized to maximize the payload capacity. Finally, the paper compares the technical characteristics of the presented winged launch vehicle to the Starship’s capabilities to highlight the key advantages of the two differing approaches and identify promising future development roadmaps for European launchers. PubDate: 2025-05-28
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Abstract: The privatisation and the commercialisation of the space sector have enabled the implementation of space missions at reduced costs, allowing small companies and universities to develop and launch their own satellites such as CubeSats, PocketQubes, and ChipSats. While CubeSats are nowadays largely used, PocketQubes and ChipSats are still emerging as platforms for space missions. This paper details the mission, system, payload design, technology experiment, and challenges of the PocketQube for In-Orbit Technology Operations, or POQUITO. In particular, POQUITO is (1) the first PocketQube mission to host an independent ChipSat onboard, (2) the first mission to test the communication link between a PocketQube and a ChipSat in visible light, and (3) the first PocketQube mission to use magnetorquers directly printed in the inner layers of the internal stack PCBs and lateral panel PCBs. The 3-axis control using printed magnetorquers can successfully detumble the PocketQube from an angular velocity of 20 $$^{\circ }$$/s in less than one orbital period, while saving volume and mass. Additionally, POQUITO adopts iterative design-development-test cycles for rapid spacecraft development and qualification, contrasting with the traditional phase-review methodology. The overall space system fits within a 5 cm edge cube and weighs 185g. Several challenges encountered in the development of the POQUITO mission, encompassing areas such as spacecraft development, launch opportunities, and non-technical matters as insurance coverage for pico-satellites are also detailed in the paper, providing useful guidelines for teams aiming to work in the sub-CubeSat area. PubDate: 2025-05-26
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Abstract: The design and manufacturing technologies used for inflatable aeroshells, a new concept in atmospheric entry systems, are being increasingly researched for improving the performance of aeroshells. In this study, we developed a feasible design method for an inflatable aeroshell to obtain a tension-shell structure with a single inflatable ring and applied it to ballistic suborbital flights of sounding rockets. The requirements of an inflatable aeroshell can be defined in terms of the maximum aerodynamic heating experienced during atmospheric entry, resistance to maximum aerodynamic loads, durability of gas filling, and terminal velocity that enables a soft landing for recovery. The upper limit of the ballistic coefficient required for the vehicle was determined based on the aforementioned requirements. This ballistic coefficient facilitated the direct determination of the aeroshell size with the largest vehicle mass considering the durability of the aeroshell. The proposed design method was used to estimate the inflatable aeroshell size required for application to sounding rocket flights; the mass of the aeroshell was approximately 20% of the total vehicle mass. Furthermore, the storage space was estimated using the storage method by wrapping a gas-evacuated inflatable aeroshell around a central pillar. The study findings confirm that the estimation is appropriate when compared with an actual flight capsule. PubDate: 2025-05-19
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Abstract: The Small Satellite Student Society at the University of Stuttgart (KSat e.V.) has a strong history of ferrofluid-based research, demonstrated through ISS projects PAPELL and FARGO, as well as the REXUS sounding rocket project FerrAS. The latest endeavor, project FINIX, is scheduled to launch on the REXUS 34 sounding rocket in Q1 2025. Building on these previously conducted missions, FINIX seeks to advance ferrofluid-based solutions with future applications in space systems. These novel developments hold the potential to replace conventional mechanical components in space applications, which are subject to wear and tear and hence represent a limiting factor for space missions. This paper details the experiments planned for the technology demonstrator mission FINIX, focusing on the development of an electrical switch and a pump, and their connection to earlier developments. The design, testing, and validation processes of these ferrofluid-based experiments will be discussed, highlighting the scientific objectives and innovations of the project. PubDate: 2025-05-17
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Abstract: In the framework of the European project MOREandLess, under Horizon 2020 program, this work aims to contribute to the growing research on supersonic aircraft for commercial use. Through collaboration between ISL and CIRA, a series of outdoor investigation activities were conducted to produce a large amount of high-quality experimental data suitable for both characterizing the sonic boom and for validating or refining existing and novel analytical methods for sonic boom prediction. The sonic boom is the massive thunder-like noise produced when an aircraft breaks through the sound barrier, representing one of the most persistent challenges of the supersonic regime. To address this issue, considerable efforts were devoted to developing various demonstrators and evaluating their impact on the intensity of sonic booms. A first test campaign, consisting of five free-flight tests, was carried out on October 4th and 5th, 2022, at the ISL firing range in Saint-Louis, France. During these tests, a reduced and slightly modified version of the STRATOFLY MR3 vehicle was launched using a 91 mm powder cannon, reaching an initial Mach number of approximately 4.7. To capture the sonic boom, ISL and CIRA deployed various types of sensors in the field. The collected data was subsequently processed to generate a sonic boom directivity diagram for the model and a comparison with Whitham’s modified linear theory predictions. PubDate: 2025-05-12
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Abstract: The Reusability Flight Experiment (ReFEx) is an experimental vehicle developed by the German Aerospace Center (DLR). It simulates the reentry of a reusable booster stage. After being launched with a VSB30 rocket to an altitude of about 139 km, it will perform an autonomous reentry. This paper focuses on the analysis of the aerothermal loads during the ascending and reentry phases. Different approaches are applied and compared for investigating thermal heating: from a very conservative worst-case analysis to a fully coupled simulation via the DLR CoNF$$^\text {2}$$aS$$^\text {2}$$ tool chain (Coupled Numerical Fluid Flight Mechanic And Structure Simulation). In the context of a numerical flight test, CoNF$$^\text {2}$$aS$$^\text {2}$$ is an integrated fluid–structure interacting coupled process chain, allowing the combination of various flow solvers, such as DLR Tau, with commercially available structural solver codes, as well as simple heating evaluations. The focus of the presented work is the heat conduction within the structure induced by the surrounding atmosphere during launch and reentry on the experimental ReFEx vehicle itself, excluding the sounding rocket motor. PubDate: 2025-05-10
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Abstract: Machine Learning (ML), or Artificial Intelligence (AI) in general, is among today’s fastest-growing methods to handle complex or computationally intensive tasks. ML is commonly implemented with Artificial Neural Networks (ANNs) on conventional computer systems that can limit their full potential. Even with access to specialized hardware such as graphics cards or Tensor Processing Units (TPUs), the demand for more computing power constantly increases. Although these hardware requirements can be met for terrestrial applications, an extraterrestrial or in-orbit application is considerably more challenging. Additional requirements for energy budget, thermal control, and radiation resistance can usually not be met, especially for small spacecraft. The benefits of an AI system for fast onboard data processing would, however, be remarkable. An optical approach to this problem can potentially be the solution. Optical computers promise to be much more energy efficient and better suitable for the mentioned space requirements. An implementation of an optical computing device on a spacecraft has not been done and can be considered as a technological leap. This work, along with the project Optical Computing for Machine Learning in Orbit (OMLO) of the Technical University of Berlin (TU-Berlin), aims to specify and conceptualize such a system. PubDate: 2025-05-02
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Abstract: This paper presents the development, characteristics, and comprehensive performance analysis of a W-band ground station receiver, designed for high-throughput satellite applications, including real-time 4K video data transmission. Operating within the 71–76 GHz frequency range, this ground station features a 57 dBi gain Cassegrain antenna system, a receiver with 5 GHz analog radio frequency bandwidth, low-noise signal reception with less than 2.5 dB noise figure, and a signal sampling rate of 16 GSa/s, with a receiver sensitivity below -130 dBm. It incorporates mechanical beam steering and monopulse multimode tracking capabilities that enable alignment and signal tracking. The station also supports autonomous operational modes and robust search patterns, essential for dynamic satellite communication environments. With the ability to display a real-time 5 GHz spectrogram and store up to 12 TB of digitized data, the station is tailored for data-intensive applications. Measurement results confirm the ground station’s superior performance in terms of signal bandwidth and receiver sensitivity, demonstrating its potential to significantly enhance satellite communication infrastructure. This work details the design of the ground station and explores its practical implications in modern aerospace electronic systems. PubDate: 2025-04-29
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