While accurate description of the internal structure of strong shock layers is beyond the scope of the Navier–Stokes equations [2, 4, 18, 30], continuum equations may still be used to describe weak shock layers at low Mach numbers. The present work extends the steady laminar base flow model of [11] and the two-dimensional modal linear stability analysis of [8] to a full three-component base flow and to two- and three-dimensional perturbations, solving the resulting eigenvalue problem by two structurally independent strategies: a matrix-forming Chebyshev collocation and an orthonormalized shooting method. The main findings are: (i) the full-spectrum-structure map, d(ω )=0 , built semi-analytically from the far-field characteristic exponents [12], locating the entire continuous spectrum over the complex frequency plane and used throughout as the reference against which the two discretization strategies are compared; (ii) no discrete eigenvalues and no unstable members of the continuous spectrum found, consistent with the results of [8] and in line with the recent analytical, Evans-function-based stability results [13]; (iii) the spectrum is proved to depend on the transverse wavenumbers k_y,k_z only through their magnitude k_t=√(k_y^2+k_z^2) , an exact SO(2) covariance of the linearized operator and an analog of Squire’s theorem, from which three-dimensional perturbations follow to be damped by exactly the same mechanism as, and always at least as strongly as, their two-dimensional counterparts; (iv) the least-damped continuous branches become less damped as the disturbance wavelength grows toward macroscopic scales, without ever crossing into instability; (v) the least-damped eigenfunctions display a damped oscillatory decay on the hot (downstream) side and, more generally, separate by branch into shock-localized acoustic disturbances and non-localized entropy/vorticity disturbances carried downstream; (vi) a general three-component steady base flow necessarily reduces, under a boundedness constraint, to the plane oblique shock, whose frequency spectrum is a rigid Doppler shift of its normal shock counterparts, damping rates remaining unchanged.
Satellites play a critical role in communication, navigation, Earth observation, and space science missions, where orbit prediction accuracy directly affects mission safety and space operations. This study proposes a hybrid convolutional neural network–long short-term memory–Transformer (CNN-LSTM-Transformer) model for future SGP4 orbit error correction under complex perturbation conditions. The model learns the temporal evolution of historical orbital features and previously observed SGP4 residuals to predict future orbit error corrections, which are then applied to the original SGP4-propagated orbit to improve prediction accuracy. Specifically, the CNN extracts local disturbance-related features, the LSTM captures short-term temporal dynamics, and the Transformer models long-range dependencies in orbital evolution. For the HY-2B satellite, the three-axis position RMSE is reduced from approximately 500 m to approximately 50 m, achieving an improvement of more than 90
Founded as a launch service provider for sounding rocket flight missions in the sixties, Mobile Rocket Base (MORABA) has conducted more than five hundred sounding rocket missions. For the longest time, focus of the research supported was on astronomy, atmospheric physics and microgravity research. Beginning in the new millennium, hypersonic research and testing has become a relevant field of engagement, spurring developmental efforts to adapt our traditional sounding rocket portfolio and flight systems to the special needs of the field. This encompassed advances in thermal hardening of exposed flight structures and suppressed trajectory designs providing flight Mach numbers up to eight for more than two minutes. Four missions also involved vehicle configurations that deviated from the traditional, rotational symmetry of sounding rockets, posing new challenges to flight stability. The present paper discusses the challenges and potential of utilizing sounding rockets in hypersonic research and presents technical adaptations demonstrated by the Mobile Rocket Base. A summary is given of the to date fourteen missions in service of hypersonic research from the perspective of flight performance and technical advances. Last, we provide an outlook on current developments aiming to meet the demand for Mach numbers in excess of ten and heavier and more complex payload designs.
As humanity continues to pursue higher flight speeds, the thermal protection requirements of aerospace vehicles increase dramatically with Mach number and flight duration. Electron Transpiration Cooling (ETC) has emerged as a highly promising active thermal protection technology for hypersonic vehicles, offering the potential to alleviate extreme aerodynamic heating through thermionic emission. This paper systematically reviews the evolution and current state of ETC research, summarizing recent progress in numerical simulations, material development, and experimental validation, while highlighting the existing limitations and challenges. Finally, several key directions for future research are proposed, emphasizing the need for accurate modeling of space-charge effects, the development and experimental verification of refractory low-work-function materials, and the establishment of synchronized temperature–current measurement techniques.
Hypervelocity impact crater formation represents a phenomenon of great scientific interest, with important applications in aerospace engineering. In this work, prediction models to estimate the size of craters resulting from high velocity and hypervelocity impacts are analysed and optimized, comparing them with the experimental data available in literature. In particular, discrepancies between predictions and observed results are highlighted and their causes are investigated; possible explanations, based on physical phenomena affecting crater formation and mathematical factors employed in the formulations, are proposed. Subsequently, a process of optimizing the prediction models is conducted, with the goal of reducing the error in crater size estimates with a custom optimized formulation. Results are presented through graphs and in terms of average deviation of the analysed models and the developed one with respect to experimental data; for normal impacts with spherical projectiles on ductile metal targets the proposed model achieves a mean relative percentage error of about 8.7
An international round-robin high-enthalpy test-facility campaign workshop was conducted during the 4th International Conference on High-Speed Vehicle Science and Technology in Tours, France, September 22nd–25th, 2025. This workshop expanded on previous discussions that resulted from a workshop held at the 3rd International Conference on High-Speed Vehicle Science and Technology in Busan, South Korea in April 2024. At this previous workshop, participants suggested that a round-robin high-enthalpy test campaign could benefit the international community by fostering collaboration and advancing knowledge of ceramic matrix composites. Building on this idea, the workshop held in conjunction with the 4th conference focused on organizing and implementing such a round-robin high-enthalpy test campaign for the benefit of the international community. This workshop brought together experts and stakeholders from leading global institutions involved in hypersonic research who utilize high-enthalpy testing to gain insights into high-temperature materials. The objective of the campaign is to organize collaboration among international high-enthalpy ground-test facilities, and advance material evaluation under extreme aerothermal environments. Primary objectives during the workshop included identifying materials for testing, defining common test geometries, aligning technical goals, and developing protocols for data sharing. A key outcome of the discussion was a need to establish a shared framework for participants, so that results can be compared and validated. This workshop kickstarted an endeavor of international cross-collaboration that will help define requirements for future experimental campaigns, while emphasizing the importance of transparent data exchange, standardized procedures, and facility‑specific characterization. These identified topics will help ensure accuracy of rapid advancements to our understanding of high‑temperature materials. An outcome of facility-specific characterization will also support a broader range of results from different high‑enthalpy facility configurations, enabling meaningful correlations from shared transparent in situ test data to potentially reduce uncertainty and future testing requirements.
The End-of-Mission operations of the Low Earth Orbit (LEO) satellite are an essential requirement and a priority for sustainable space operations. Meticulous postmission disposal activities help to minimize orbital life of the spacecraft and reduce the risk of creating additional orbital debris. Inter-Agency Space Debris Coordination Committee (IADC) in its space debris mitigation guidelines, Mar 2020, recommends for “Objects Passing through the LEO Region” that the residual orbital lifetime of a LEO spacecraft should be less than 25 years. As a part of IADC member agencies, ISRO has successfully made its two satellites MICROSAT-TD and RISAT-2 to safely re-enter after the mission life of 3 and 13 years, respectively. This paper discusses RISAT-2 spacecraft maneuver plan to extend its payload operations for the maximum usage and re-entry predictions from 12/Oct/2022 to 29/Oct/2022. The in-house developed high-fidelity numerical orbit propagator known as Satellite Precise Orbit Propagator (SPOP) is used to precisely predict the re-entry time and location of RISAT-2. Further, the re-entry predictions using the STK Lifetime option, the commercial package of Ansys/Synopsys and with Space-Track.org are carried out for the comparative study with respect to SPOP. The observed orbit decay of the spacecraft at various altitudes is also described.
This study explores the development of multifunctional sandwich panels for spacecraft, focusing on their effectiveness against both micrometeoroid and orbital debris (MMOD) impacts and extreme space radiation. The research incorporates advancements from the University of Manitoba Orbital Debris Research (UM-ORDER) group and the Composite Materials and Structures Research groups. It integrates the aluminum foam-core sandwich panels (with improved ballistic performance) with the radiation shielding benefits of polyethylene and carbon-fiber-reinforced composite, particularly in high-radiation Molniya orbits. Seven distinct panel designs were examined, each featuring a modified MMOD/MLI bumper design and varying facesheet materials like aluminum, polyethylene, and carbon-fiber-reinforced composite. The designs were numerically analyzed for radiation and ballistic protection and compared against traditional aluminum panels of equal weight. The results indicated enhanced radiation protection in polyethylene and carbon-reinforced composite panels, and except for one model, all designs showed effective ballistic protection without rear-facesheet perforation.
Research on the radiative forcing and climate impact of emissions from the space transportation sector remains in its early stages. A critical component of advancing this field is the development of accurate global emission inventories, which depend fundamentally on the precise characterization of rocket engine exhaust products. One major current limitation is the reliance on simplified exhaust datasets that fail to capture the chemical and mass flow complexities introduced by different engine turbopump cycles. To address this gap, this study establishes a comprehensive new reference dataset of primary rocket engine emissions (modeled up to the nozzle exit plane). The dataset is anchored by a consistent 2200 kN thrust-class baseline, encompassing major cryogenic propellants, a hypergolic combination, and a solid rocket motor. For liquid propellant systems, both closed cycle (staged combustion) and open-cycle (gas generator) architectures are evaluated. Additionally, an exploratory methodology is introduced to assess the impacts of fuel film cooling by modeling a fraction of the total mass flow at a distinctly fuel-rich oxidizer-to-fuel ratio. Both the specific engine cycle and the implementation of film cooling are demonstrated to be critical factors in predicting black carbon exhaust fractions. Finally, these newly established reference datasets are compared against existing inventories. By employing 0D equilibrium combustion calculations at realistic oxidizer-to-fuel ratios and explicitly accounting for the distinct mass flows of different turbopump cycles, this study provides a highly consistent exhaust dataset to improve the fidelity of future atmospheric and climate models.
The performance of a geolocation system consisting of three collectors: one in a high-eccentricity Molniya orbit (MEO) and two in low-inclination geosynchronous orbits (IGEO) is evaluated. Bias error sources for the collectors are quantified: orbit determination accuracy; tropospheric path delays; ionospheric delays and frequency shifts; relativistic Doppler shift; light-transit-time (LTT) delays in the signal reaching the collector; and the integration time used to process the communications signal and obtain time- and frequency-difference-of-arrival measurements. An analytical approach to mitigating the impact of these biases on geolocation data is presented. Monte Carlo (MC) and covariance analyses are used to validate the models and assess the performance of the conceptual geolocation system, using the mean geolocation position and containment as metrics. Within ± 2 h of the Molniya apogee, a Molniya plus 2 IGEO (M2G) constellation can provide a performance advantage over a 3-IGEO (3G) constellation twice daily. However, this advantage is highly sensitive to Molniya orbit determination errors.
In recent years, electric propulsion has gained increasing importance in space applications driven by the growing availability of onboard electrical power in modern spacecraft. Among the various concepts, Hall thrusters stand out for their relatively high specific impulse, thrust capabilities, and scalability, while magnetic shielding significantly enhances their operational lifetime. An alternative variant is the Thruster with Anode Layer (TAL), which not only offers improved durability but also allows for a more compact design compared to conventional Hall thrusters. In addition, the use of alternative propellants, such as Argon, is being explored to reduce operational costs and increase propellant availability. This work focuses on the scaling of TAL Hall thruster operating at approximately 100 W power level. A scaling methodology based on a phenomenological model was applied to derive two candidate design configurations. The approach was assessed through comparison with experimental data from an available TAL thruster performance database. The comparison shows trends consistent with the reported experimental performance parameters. The resulting designs were further compared with existing TAL thrusters operating at similar power levels, providing comparable estimates in terms of efficiency, specific impulse, and geometric compactness. These results support the applicability of the proposed methodology for preliminary design studies of low-power TAL thrusters and provide a reference framework for future experimental validation. The present study represents the first phase of a broader development effort, which also includes the preliminary design of the two thruster configurations supported by magnetic, thermal and mechanical simulations. Within this subsequent phase, specific design features were considered to potentially facilitate operation with Argon.
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.
This work experimentally characterizes the cold-flow field of a clustered, truncated linear aerospike nozzle with and without outer flow. Planar particle image velocimetry and time-resolved schlieren measurements are performed in a controlled wind tunnel environment. Two inter-module spacings are realized by means of a fully activated cluster and a layout with alternately deactivated modules. In the case of nozzle flow only, increased spacing splits the composite plume into largely uncorrelated plumes, each with a narrower shock cell system. The outer plumes expand laterally into the ambient, while the center plume experiences higher effective back pressure, resulting in shorter shock cells. Two unsteady modes appear across the nozzle pressure regimes: a plug base wake mode (shear layer–trailing shock interaction) and a jet screeching mode. When increasing the inter-module spacing, jet screeching is recorded to be suppressed, and two slightly offset plug base wake mode frequencies are observed. With outer flow, the jet unsteadiness increasingly localizes toward the plug base and shock interaction regions. The measured flow field is dominated by outer flow modes (bluff-body vortex shedding at sub/transonic, shear layer flapping at supersonic outer flow). No nozzle flow modes are detected within the field of view when the outer flow is present.
We discuss the design of a hold-down and release mechanism for loads exceeding 10kN that locks up to six degrees of freedom, fully retracts using only linear motion, reduces the risk of fretting, and is simple in mechanical complexity. This makes our solution especially interesting for ultra-high-precision pointing applications, such as the optical angular tracking mechanism for LISA.
An important aspect influencing the economic advantage of reusable launch vehicles over their disposable counterparts is the cost associated with the reuse itself. These costs are optimized through design for inspection and maintenance, as well as effective condition monitoring and predictive maintenance. In this context, the fouling of rocket cooling channels due to the thermal decomposition of hydrocarbon fuels, known as pyrolysis, is particularly challenging as cooling channels are difficult to inspect and clean. Carbon depositions resulting from the pyrolysis process act as thermal insulation, due to which wall temperatures can rise and the thermo-mechanical damage to the wall can increase. The current work proposes an approach to the monitoring of cooling channel fouling based on the sensing of decomposition products in the fuel flow. In addition to acting in real-time, this approach provides an early warning of damage that will occur, as it is based on reaction products that precede the deposition of carbon. It is inherently faster than the alternative approach of sensing increases in the wall temperature. The decomposition products are sensed by thermal conductivity gauges and, using a simplified theoretical model, the data is used to infer the mass of deposited carbon in the channel. The approach is tested experimentally for methane-based fuels on Inconel 600, Inconel 625 and Haynes 230. The model is shown to match carbon deposition mass measurements to within approximately the uncertainty of the scale used. Although these results are promising, further testing under more arduous conditions needs to be conducted in future work.
Compliant mechanisms achieve motion through elastic deformation rather than contact surfaces, eliminating friction, wear, and lubrication requirements. Historically, conventional manufacturing processes have constrained both the design and adoption of these mechanisms in cryogenic and space applications. This work exploits the design freedom of Laser Powder Bed Fusion (LPBF) to propose a streamlined Design for Additive Manufacturing (DfAM) workflow for optimizing and customizing a flexure pivot operating at 4.2 K within the Mode Selector Mechanism (MSM). The methodology comprises three stages: (i) material selection and post-processing—316L stainless steel combined with stress annealing and HIP—to ensure cryogenic compatibility and fatigue strength; (ii) optimization of the Interlocked Lattice Flexure (ILF) geometry considering performance metrics (stroke, stiffness, guiding accuracy) and AM constraints (overhang, minimum feature size); and (iii) flexure thickness tuning to meet project-specific requirements while mitigating AM-induced variability, enabling rapid customization. Numerical and experimental validation confirmed compliance with design targets: ± 3.5^∘ stroke, stiffness of 3.97 mN m/ ^∘ , and fatigue life exceeding 10^6 cycles under over-testing conditions (115
In planetary satellite systems, identifying low-energy, stable orbits that reduce interplanetary mission costs are critical for exploring minor celestial bodies in a variety of highly complex gravitational environments. To identify suitable trajectories in planetary systems with multiple moons, prominent ring systems and varying levels of planetary oblateness, the application of the Circular Restricted Three-Body Problem (CR3BP) framework is proposed. Classical celestial mechanics demonstrates that CR3BP may generate an infinite number of periodic orbits. Nonetheless, the original framework represents an idealized scenario that does not consider the dynamic influences of perturbations, such as those that stem from planetary rings and the oblateness of celestial bodies, leading to incomplete assessments of orbital stability. The proposed CR3BP-based framework involves the implementation of a perturbed propagator that can automate the generation of large periodic orbit databases via pseudo arc-length continuation (PAL). In this context, the Saturn–Titan system is employed as a relevant case study to demonstrate the capability of this framework to simulate complex gravitational dynamics. Furthermore, the behavior of Distant Retrograde Orbits (DROs), Southern and Northern Halo orbits around one of the collinear equilibrium points in this planetary system is also analyzed to demonstrate the applicability of this model to multi-body environments. Results indicate that despite the introduction of these perturbations, the model is capable of capturing low-energy trajectories that may offer significant potential for designing sustainable, more cost-efficient deep-space missions in the Saturn-Titan system. The proposed approach is applicable to any planetary system with similar characteristics, thereby providing a foundation that may be extrapolated to higher fidelity models for the design of future interplanetary space missions.
Spectral markers are materials with a characteristic emission spectrum, which are added to a spacecraft structure, e.g., a satellite. During an atmospheric re-entry, the marker is released and its radiation can be detected by appropriate instruments. This allows to identify breakup events and to assess the spacecraft’s demise behavior during an observation campaign. In this paper, we study various compounds to select suitable materials for future integration as spectral markers. Six material candidates were tested in the plasma wind tunnel PWK4 at the Institute of Space Systems at the University of Stuttgart. Powdered material was encased in an epoxy resin matrix and mounted to a water-cooled sample holder. Spectra, temperature, and visual data were recorded throughout testing. Lanthanum acetate, lanthanum oxide, and lanthanum oxalate are considered non-viable marker materials. Lanthanum hexaboride features strong atomic lines of lanthanum as well as molecular bands of LaO and BO in a wide wavelength range within the visible and near-infrared (370–820 nm). Rubidium chloride features very strong atomic lines of Rb in the NIR (780–790 nm), well separated from the atomic radiation of air species. Strontium features moderately strong atomic lines. In conclusion, rubidium chloride and lanthanum hexaboride are considered promising marker materials for re-entry event identification through spectroscopic observation.
The combustion characteristics of tri-propellant coaxial shear injectors with Air/LOX/ethanol propellants, featuring different outlet geometries, were investigated through hot fire experiments conducted using a combustion air heater. The pressure time evolution and high-speed camera images were acquired for all test cases, enabling comparisons of the spray morphology, ignition process, and flame dynamics across different injector geometries. Analysis of the spray process revealed that the expanded outlet structure introduces an impingement mode within the coaxial shear injector, enhancing liquid-liquid interaction and thereby improving atomization and promoting LOX evaporation. The ignition process of the combustion air heater was found to comprise two distinct stages: formation of a flame core downstream of the liquid jet, followed by upstream propagation of the flame core. Furthermore, grayscale intensity analysis revealed periodic flame variation closely corresponding to the chamber pressure time evolution, indicating that the instantaneous heat release coupled with the chamber pressure fluctuations, thereby amplifying the pressure oscillations. The injector base height and chamfer half-angle were identified as parameters significantly governing ignition and combustion characteristics, with the chamfer half exerting a great influence. Among the configurations investigated, a 15 ^∘ chamfer half-angle proved most favorable for achieving reliable ignition and maintaining stable combustion.
Accurate prediction of ionization in hypersonic reentry flows is essential for assessing plasma effects on electromagnetic wave propagation, including radio blackout and radar signature alteration. This work investigates the sensitivity of electron number density predictions for the RAM-C II flight experiment using the in-house CFD solver NExT. The simulations solve the compressible Navier–Stokes equations for an 11-species, chemically reacting air mixture in thermal and chemical non-equilibrium, employing a multi-temperature formulation for vibrational energy. A systematic sensitivity analysis is then conducted to assess the influence of chemical kinetic schemes, equilibrium constant formulations, vibration–chemistry coupling, wall catalysis, and three-dimensional angle-of-attack effects. The results indicate that the Park’85 mechanism systematically overpredicts electron density, whereas Park’93 and Kim show closer agreement with flight data. Incorporating fitted equilibrium constants and vibration–chemistry coupling significantly improves the predictions. Wall catalyticity also plays a major role, with a species-selective treatment—non-catalytic for neutral species and catalytic for charged species—yielding the best overall agreement with measurements. Three-dimensional simulations further reveal that small variations in angle of attack can strongly modify the local electron density through windward compression and leeward expansion; however, these effects weaken at higher altitudes, where diffusion becomes dominant. Overall, the study quantifies the impact of key thermochemical modeling assumptions on electron density predictions in hypersonic reentry flows and supports the capability of NExT to accurately capture weakly ionized plasma conditions.
