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This workshop is organized by the Working Group “ATD3”, managed by ESA and CNES through the ESA Technology Directorate and the CNES Research and Technology Directorate.
As in all the previous editions, this workshop will be an opportunity for research centres, universities, industries, and space agencies to exchange ideas, advance knowledge and discuss about the following topics:
The format of the event will be a sequential list of presentations of about 15 minutes each, with 5 minutes brief Q&A session per presentation.
WORKSHOP REGISTRATION: CLOSED
The participation to the workshop is free of charge, but the online registration is required. The networking dinner is not included.
If you want to join us for the dinner at the ZOMERS Beachlub, in Katwijk ann Zee, on June 6th 2024, please register and pay your fee here:
NETWORKING DINNER REGISTRATION
Ticket: 45 euros (drinks excluded)
CALL FOR ABSTRACT: CLOSED
Please note the following deadlines:
31 May 2024: Extended dinner registration deadline
20 May 2024: Registration deadline for the workshop
29 April 2024: Extended abstract submission deadline
3 May 2024: Abstract selection announcement
24 May 2024: Complete program online
6-7 June 2024: Workshop
THE ATD3 WORKING GROUP
The ATD3 Working Group aims to achieve the following objectives:
At the end of their lifespan, uncontrolled spacecraft eventually re-enter Earth’s atmosphere and demise, totally or partially, due to their interaction with the surrounding flow. The growing population of junk in Earth’s orbits induced space agencies to tackle the space debris issue by imposing increasingly stringent requirements over the years. The recently published standard ESA Space Debris Mitigation Requirements ESSB-ST-U-007 recommends Design for Demise (D4D) as the best choice among the possible end-of-life mitigation strategies that can be adopted to limit the risk of generating casualties on the ground. The D4D philosophy strongly advocates that, early in the design and development stages of a new space mission, careful evaluations of spacecraft buses, payloads, and structural components should be conducted to assess their potential to survive an uncontrolled re-entry.
Currently, the demise of spacecraft is assessed using low-fidelity engineering codes (object or spacecraft-oriented) that enable the simulation of the degradation of the satellite along the entry trajectory. Those engineering tools rely on strong hypotheses about heat flux prediction, ablation phenomena or structural failure of the spacecraft. Within the context of the ESA project DSMCFED (contract no. 4000135337/21/NL/MG), we are developing a high-fidelity design toolset to complement existing engineering tools. The methodology enables a more comprehensive understanding of spacecraft fragmentation through detailed modelling of the key aerothermodynamic, thermal, and structural phenomena contributing to such events. This is achieved by coupling high-fidelity physics-based numerical tools able to simulate the spacecraft aerothermodynamics in the rarefied/transitional regime and the thermo-structural behaviour of the spacecraft structure.
The SMURFS (Spacecraft Motion and behaviour Under Re-entry for Fragmentation Simulations) toolset integrates trajectory, flow, and thermo-mechanical computations. Based on a volumetric mesh of the spacecraft, the decomposition of the geometry into fragments during the trajectory is tackled by the toolset. The methodology employs a loosely coupled approach among the three modules. Steady-state aerothermodynamics evaluations at predefined altitude stations are combined with transient 6-DoF (Degrees of Freedom) trajectory analysis and quasi-static thermo-mechanical responses. This setup allows for the fragmentation assessment through predefined failure criteria, identifying potential breakup events with dedicated thermal and mechanical criteria. Thermal criteria are used to estimate fragmentation when the melting temperature is reached, and the mechanical properties are strongly degraded. Mechanical failure criteria are used in the second simulation step to detect fragmentation due to stress states. When fragmentation occurs, each resulting debris becomes independent from the others and is managed as a new simulation branch to compute further decomposition.
In this presentation, we will provide an update on the status of our developed toolset, the progress of ongoing initiatives, and the application to various test cases. We will outline the toolset's features, discuss the foundational assumptions, and offer guidelines for conducting simulations.
The French Space Operation Act (LOS) adopted in 2008 has established a national regime of authorization and supervision for space activities. CNES, the French Space Agency, is in charge of ensuring the right application of this Space Operation Act. In this context, to predict the debris survivability of a space vehicle and its associated fragments during their atmospheric re-entry, and assess the prospective risk on the ground, CNES develops its own spacecraft-oriented tool named PAMPERO. PAMPERO is a multidisciplinary tool that models the complete atmospheric reentry of a spacecraft including the fragmentation and ablation processes, along a six DOF trajectory.
Over the past 30 years, numerous methods and tools have been developed to simulate spacecraft breakup during atmospheric re-entry, predict the characteristics of the surviving fragments, and estimate the ground casualty risk. With the introduction of the Design for Demise concept, these tools are used for designing spacecraft that eventually break up and burn up during re-entry, thus usually reducing debris impact risk. To enhance both the accuracy and efficiency of predictions, researchers made continuous improvements in this field, especially in the last decade, but uncertainties and gaps in knowledge remain.
The PAMPERO Hifi (high-fidelity) code has been developed since 2019, with the aim of using high-fidelity codes to deal with the physics of atmospheric re-entry. Recent studies show that the parameters that have the most impact on the spacecraft's demise are the predictions of the heat load and the fragmentation process. It is therefore of prime importance to close the gaps that are existing in the spacecraft-oriented codes,
Especially :
Start discarding correlations based on Modified Newton laws, as the shock interaction and wake flows cannot be accounted for. New methodology has been developed based on CFD/DSMC computations and Machine Learning Techniques : For continuous regime : activity ongoing between 2019-2024; For rarefied regime activity ongoing between 2024 and 2026
Take into account the detailed geometry and the aerodynamics load at each time step to compute a mechanical analysis. Activity completed between 2019 – 2022
Take into account the in-depth Pyrolysis process and gas fluxes generation, Mass transfer fluxes, Advection fluxes,Char creation, char ablation process and associated fluxes with outgasing phenomena activity ongoing between 2021-2024
The aim of this presentation is to show the strategy implemented in PAMPERO Hifi in order to be able to estimate the aerothermodynamics load and the fragmentation/ablation process with precision, while keeping computation time to a minimum.
The risk of fragments surviving a destructive re-entry and impacting on ground, the state of the art around the turn of the millennium was based on thermal protection shield assumptions and limited material ablation testing. A significant amount of research in this field during the last decade indicated that the simplifying assumptions on shape and complex geometries lead to uncertainties of 50-100% and above when it comes the heating processes alone, and hence have a driving influence on the physical predictions of the demise behaviour. On-ground wind tunnel testing on spacecraft component indicated a strong dependency on flow based length-scale, i.e. local and global transition regimes, which even the most advance system level tools don’t account for.
The "Shape Effect Modelling for Risk Evaluation" project is dedicated to improving risk models with a critical dependency on shape assumption, and is embedded in the Agency’s Space Safety Programme. The general objective of the Space Safety Programme is to contribute to the protection of our planet, humanity and assets in space and on Earth from hazards originating in Space and to contribute to Europe, by providing safety from such hazards, as a service to its society. This activity focusses on analysing the probability of a hazard materialising and establishing its severity and magnitude; part of the Space Debris track of the Core Activities of the programme.
The top-level objective of the re-entry track of this activity is to increase the accuracy and the precision of the space debris risk methodologies in the field of destructive re-entry break-up modelling. This shall be done by improving the underlying physical response models used by such risk assessment methodology, rather than the methodologies that combine the likelihood of a physical response taking place with the severity of this event. E.g. it is envisaged to improve the aerothermodynamics coefficients in the transition regimes, but not the estimations of casualty areas.
The top-level objectives of the re-entry part of the activity are:
- Identify and assess the impact of shape assumptions, understood as the determination of characterising length scales and complex shape and behaviour in the transition regime, on the demise and fragmentation of spacecraft structures during destructive atmospheric re-entry.
- Develop a software library to implement the improved physics models into risk assessment methodologies.
These objectives shall be achieved by developing an "Aerodynamic/Aerothermodynamic Shape Effect Override (ASEO)" prototype. The ASEO prototype shall provide aerodynamic/aerothermodynamic coefficients derived from wind-tunnel experiments or high-fidelity CFD simulations for specific shapes.
At enGits, we have developed an in-house CFD code for compressible flow. From the start, this code has been designed with massively parallel hardware in mind. As a result the code is able to run very large simulations on simple PCs or similar hardware. For example, we are able to run a buffeting simulation with 100 million cells overnight on a single PC. We believe that this is interesting for demise simulations, because a large number of different conditions and configurations will need to be simulated.
The software was initially developed using immersed boundaries on Cartesian grids, which allows to effectively run many different geometries at very little pre-processing costs. Unfortunately this has often led to unwanted modeling compromises, because a detailed resolution of boundary layers with the associated friction and heat exchange effects was not always possible.
Over the past year, the code was extended in order to be able to handle arbitrary grids. Thus we are now able to run simulations with resolved boundary layers while still having a similar efficiency compared to running on Cartesian grids alone. The software can run on purely Cartesian grids, or purely unstructured grids, or on a mixture of both grids types. For the later hybrid approach we use an interpolation method similar to what is used in the overset mesh approach.
Considering the market of CFD for space applications, the next logical approach for the software development is to extend it to chemically non-equilibrium and equilibrium modeling for re-entry simulations. We have already been approached by potential customers who are interested in such a capability and we are keen to implement these features into our software in order to fill a vital need of the emerging new space industry in Europe. As of today, there are not many commercial CFD packages available which can tackle this kind of problem.
Furthermore we believe that this would also be interesting for satellite and spacecraft demise simulations. Hence we would like to present the current state of the software at this workshop and show what could already be done with it in the field of spacecraft demise. Additionally we are also seeking potential partners to bring this development forward and believe that the workshop could be a good platform to present the current state and to find ideas and possibly partners for future developments.
Atmospheric re-entry tools and their predictions are increasingly utilized in the satellite design process due to the demand for safe satellite disposal to minimize the rising debris population in orbit. Accurate predictions of re-entry, particularly in aerothermodynamics, are crucial to evaluate the probability of demise and casualty risk associated with re-entering vehicles. Destructive atmospheric re-entry is an inherently complex environment to model, with hypersonic re-entry eliciting shock waves, thermal and chemical non-equilibrium and high-temperature flows. Traditionally, aerothermodynamics are predicted using low- or high-fidelity methods. Low-fidelity methods, based on semi-empirical and analytical correlations are computationally efficient but disregard the complex physics of destructive re-entry. This leads to substantial uncertainties in the predicted fragmentation and trajectory of the vehicle. Conversely, high-fidelity methods such as Computational Fluid Dynamics (CFD), are increasingly capable of resolving the complex phenomena of re-entry, but at a significantly high computational cost. Achieving a balance between the accuracy and cost of re-entry predictions is crucial for the design for demise processes of satellites.
Emerging research explores a promising approach using Deep Learning techniques to create a synergy between low- and high-fidelity methods to establish the balance of accuracy and computational efficiency. Deep learning methods, such as the neural network architectures of feed-forward artificial neural networks (ANN) and convolutional neural networks (CNN) leverage a set of high-fidelity simulations to predict surface quantities distributions like heat-flux or pressure. These predictions incorporate the complex physics of numerical high-fidelity simulations to provide a superior accuracy when compared with low-fidelity or physically-blind methods. The computational cost of a given prediction by deep learning models is however computationally inexpensive, allowing for the probabilistic analysis of design for demise processes of satellites to be carried in a computationally efficient and accurate capacity.
The purpose of this study is to investigate the implementation and applicability of deep learning models for the prediction of aerothermodynamic loads during destructive hypersonic re-entry. Geometrically simple test cases, such as the ring and cube will be considered to analyse the training data demand of deep learning models, as well as their ability to resolve complex non-linearities of re-entry for rarefied and continuum flow. The Automated Transfer Vehicle (ATV) will also be used to investigate the deep learning models applicability to complex test cases in the continuum flow regime. The optimal neural network architecture and training procedure for these types of problems will be explored as part of this work. Python’s machine learning library TensorFlow will be used for the creation and deployment of the neural network models of this study. In addition, the multi-fidelity re-entry tool TITAN will be used for all time-propagate re-entry analysis and the high-fidelity tools SU2-NEMO and SPARTA are utilized for high-fidelity simulations in the continuum and rarefied flow regimes, respectively. This study will therefore demonstrate the potential feasibility, and subsequent impact during atmospheric re-entry simulations on satellite trajectory and demisability when deep learning neural network models are employed. The findings will be assessed in terms of the accuracy and computational efficiency delivered by these models.
As the number of objects in LEO increases, Design for Demise (D4D) becomes an ever-more important philosophy for space engineering which relies upon robust modelling of re-entry processes. Uncertainty Quantification (UQ) is a vital part of ensuring this robustness but is limited by the significant computational costs associated with high-fidelity re-entry simulation and as such most UQ approaches only consider low-fidelity object-oriented codes.Recent developments in information fusion for Monte Carlo estimators enable multifidelity-based methods for UQ which have significant potential for D4D applications. Estimator algorithms have been developed which can leverage the correlation between high and low fidelity models to enhance the convergence rate of the estimator through variance reduction. In this work the multifidelity capabilities of the TransatmospherIc flighT simulAtioN tool (TITAN) are used as models within the estimator framework developed by Schaden and Ullman, named the Multi-Level Best Linear Unbiased Estimator (MLBLUE), in order to perform UQ on a re-entry test case. The open-source MLBLUE implementation and extension developed by Croci, Willcox and Wright, BLUEST, enables the automatic selection of models from the multiple fidelity options provided by TITAN and allocation of samples in order to minimise error for a given computational budget via Semi-Definite Programming. This enables computationally feasible estimation of the mean values of landing location whilst still incorporating high fidelity information. This results in an optimal convergence rate with increasing budget. The test case provides initial conditions derived from a Two/Three Line Element file which have inherent uncertainty in translational information. This uncertainty is propagated by the MLBLUE estimator into a landing distribution. This distribution can be used to better quantify the performance of deterministic simulation processes by using statistical distance metrics in addition to the broader contexts of application in risk analysis in the definitions of declared and safe re-entry areas.
DRAMA (Debris Risk Assessment and Mitigation Analysis) is a comprehensive tool for the compliance analysis of a space mission with space debris mitigation standards. As part of the suite DRAMA allows modelling of the re-entry of a spacecraft into the Earth’s atmosphere along with an assessment of ground casualty risk.
Presented herein is an overview of two tools, Event Simulator and PRODUCERS which take DRAMA output in order to perform additional analysis via post-processing. Both tools are Python toolsets which break down the analysis into a series of tasks.
The Event Simulator models an “Instrument” embedded on the re-entering vehicle. The instrument aims to collect data on the demise of the spacecraft and is developed within the context of the ESA DRACO program. The Event Simulator aims to provide a digital model of the flight model to aid the setup of the flight instrument prior to flight. Shown is how additional contextual data, such as connectivity, in-depth temperature, pressure and strain measurements can be extracted via post-processing. Such data is useful when mapping across to ground tests. Also shown is how contextual image data is rendered in order to replicate the response of what embedded cameras on the re-entry vehicle will see during re-entry. This includes effects due to temperature and fragmentation.
PRODUCERS, a tool for Predicting the Spectrographic Response of Break-Up Fragments to the Re-Entry Environment. PRODUCERS aims to take the first steps towards providing a capability to compare destructive entry tool predictions with spectroscopic measurements obtained in remote observation campaigns. The intent is to provide evidence to instrument designers and observation campaign planners regarding the spectral content and intensity of spectroscopic signals from destructive entry. The radiation sources PRODUCERS considers are: grey body radiation from fragments, shock layer radiation around fragments, gas phase ablation products from fragments, grey body radiation from hot particles produced by mechanical ablation of fragments, and gas phase ablation products from stripped particles. The long-term aim of PRODUCERS is to improve the ability to correlate modelled fragmentation events to observed flight spectra thus providing indirect evidence to validate these models.
Both the Event Simulator, DRACO as a whole and PRODUCERS have a main goal of providing evidence to improve the understanding of destructive entry fragmentation events. The methodology used in simulating fragmentation is a significant, if not the main, driver in assessing the ground casualty risk of a destructive entry, which strongly motivates the pursuit of these endeavours.
Dynamic fragmentation is a process during which a material or structure subjected to intense loads fails catastrophically through the initiation, propagation and coalescence of a multitude of cracks. It is a key topic in many fields of engineering, as for instance in aerospace industry, where the outcome of destructive re-entry is of great concern. Robust numerical models are direly needed to develop a fundamental understanding of such events, in particular to be able to predict the statistical distributions of fragment sizes, shapes and velocities resulting from destructive events.
A well established way to address this problem is to use finite elements solid mechanics models coupled with cohesive elements [1]. Cohesive cracks give an explicit representation of crack surfaces and simplify the treatment of contacts between fragments, a crucial factor to predict debris velocities. However, the cohesive approach is known to suffer from mesh dependency, with crack paths that depend on the underlying mesh, resulting in non-robust predictions of fragments shapes.
Phase-field modelling of fracture belongs to another family of methods using diffuse crack approaches and, as opposed to cohesive models, where the fracture paths are not dependent on the underlying mesh. Phase-field has been shown to lead to promising results in many problems, not only in quasi-static but also in dynamics where different mechanisms such as branching can be observed [2]. In addition, the multiplicity of possible crack patterns obtained with phase-field for a single loading case in quasi-static has been discussed in [3] where different simulation outcomes can be associated with probabilities. This study is yet to be extended to dynamic cases, where the evolution of statistics on fracture paths with the loading rate can be explored. This stochastic approach is of great interest in the context of dynamic fragmentation to enrich statistical data of fragment sizes and shapes in light of material heterogeneity.
The phase-field approach to fracture, addressing the issue of mesh-dependency, will be examined in fragmentation dynamics and numerical data shall be compared to available analytical models and to the outcomes of cohesive crack models. In addition, the sensitivity of debris distributions to small variations in the material parameters or the geometry of the model will be analysed.
In recent times, satellite systems boasting thousands of satellites have been launched into orbit. The relatively short lifespan of these satellites, typically lasting only 3-5 years, has lead to increasing chances of in-orbit collisions and overall space cluttering. As the number of space debris experiencing post-mission uncontrolled re-entry rapidly increases, there is growing concern for the threat of ground impact and collisions with operating spacecraft. The complete and controlled demise of spacecraft that have reached the end of their lifespan has therefore risen to the forefront as a matter of utmost importance in sustainable space exploration.
This pressing scenario underlines the crucial need for a specialized, sustained monitoring detection system designed to identify and characterize re-entry events from orbit. The information gathered from these observations, coupled with rigorous processing and analysis, has the potential to greatly improve our comprehension of re-entry phenomena. Moreover, it can contribute to refining re-entry models, facilitating
precise predictions regarding the timing and location of spacecraft re-entries and associated demise.
The work proposed in this abstract is part of the ESA project "Detection of IR to UV Re-Entry Signatures from Orbit" that has as main objective to design a wide-band orbital detector system to record hundreds of kilometre long streaks emitted from the Earth's atmosphere due to destructive re-entries of space debris in the IR, Visual, and UV spectrum. One of the main tasks of the project is the simulation of destructive atmospheric re-entry and associated radiative spectra emitted by the fragments, for the purpose of validating the detector system design.
The simulation of destructive atmospheric re-entry events is performed with the TITAN (TransatmospherIc flighT simulAtioN) code, a multi-fidelity tool that has been developed in the context of previous ESA projects. This multi-disciplinary framework combines low- and high-fidelity aerothermodynamics, thermal analysis, flight dynamics, and structural analysis in a modular approach to achieve a favourable trade-off between cost and accuracy. The present abstract introduces the most recent developments in TITAN that focus on improving the accuracy and robustness of the thermal modelling by coupling TITAN to PATO (Porous material Analysis Toolbox based on OpenFoam). PATO is a modular analysis platform for multiphase porous reactive materials, but it can be run as a simple Fourier heat transfer code. In the coupling methodology, for each trajectory point, TITAN provides information on the shape and temperature of the fragments, as well as surface convective and radiative heating, to the PATO software. Detailed heat transfer models accounting for gas-surface interaction processes are then used inside PATO to retrieve the new temperature of the fragments and the chemical species emitted during ablation. The improved modelling of the surface temperature distribution of the fragments and the information on the emitted chemical species will be the base of the emission spectra modelling to be carried out at a later stage of the project.
Direct numerical simulations (DNS) are performed by our in-house flow solver INCA (www.inca-cfd.com)(Başkaya et al., Computers & Fluids, 2024) over a 15 degree compression ramp undergoing ablation at Mach 8. The setup is validated against experiments and simulations that considered the laminar flow over an inert ramp. Streamwise vortices, which generate heating-cooling streaks through the lift-up effect, are introduced by perturbing the base flow. The ramp is then replaced by a low-temperature ablator, camphor, in our DNS and the interaction of the streaks with the recessing ablator surface is examined. We present the first findings regarding the influence of ablation on the perturbation evolution and transition to turbulence for this configuration. Effect of the ablation products being blown into the boundary layer, the recession speed-up factor, the type of surface balances employed, and the recession patterns are discussed. We investigate the destabilization of the flow due to recession and further investigate pattern formations for higher amplitude perturbations.
One of the peculiarities of some space debris materials compared to thermal protection systems is that the ablation of the critical parts of such objects might involve melting and the presence of three phases (gas, liquid, solid). This is the case for glassy and metallic materials. Given the huge properties differences that exist between the gas and the condensed phases and the complex interactions occurring at the interface, the study of this kind of situation is particularly challenging from a numerical point of view.
To complement the available low-fidelity engineering tools, we have been developing at Cenaero two main strategies inside our in-house high-order multiphysics platform, named Argo, based on the discontinuous Galerkin spatial discretization of the governing equations. The first one relies on a loosely-coupled approach in which the gas and the material solvers exchange information across a boundary in a staggered manner. The second approach considers a strongly coupled formulation of the problem by means of a monolithic sharp interface solver in which interface jumps conditions are enforced.
In the framework of the GSTP activity “Validation of Space Debris Demise Tools Using Plasma Wind Tunnel Testing and Numerical Tools” (led by the von Karman Institute), we applied the staggered approach to reproduce ablation experiments on silicate (quartz and Zerodur®) and titanium samples, conducted in the Plasmatron facility of the von Karman Institute. The multispecies reactive gas solver provides to the material solver the shear stress, the heat flux and the pressure at steady state. In the energy balance at the interface, evaporation and re-radiation are also taken into account. The mass balance is exploited to compute the recession velocity of the interface, which is treated as a moving immersed boundary (i.e. unfitted to the underlying mesh). Inside the material volume, enthalpy absorption during the melting process is modeled and the velocity from the liquid to the solid regions is penalized by a Darcy term. On the other hand, the material solver gives the interface temperature to the fluid solver that iterates until steady-state is reached. Exchanges between both solvers are performed at predefined times based on the expected material response.
Comparisons to experimental results revealed that catalytic effects must also be taken into account for silicate materials while passive and active oxidation dominate for titanium in air environment. Finally, shock capturing capabilities have been added to the code such that experiments in supersonic conditions could also be reproduced.
This presentation will cover the modeling and numerical aspects of the staggered and monolithic strategies as well as their application to experimental configurations.
DRACO (Destructive Re-entry Assessment Container Object) ESA mission will be the world’s first demonstration of a controlled break-up process of a satellite during re-entry. It is based on a representative small satellite platform with surviving capsule to transmit data on ground.
DRACO mission objectives are three-fold: to demonstrate the break-up process of a spacecraft during re-entry to extrapolate ground-test to flight, to establish an understanding by recording the physics of destructive aerothermal break-ups not accessible from ground or by model, and to test early fragmentation design for demise (D4D) technologies.
The presentation will cover the current development status of DRACO mission, together with the scientific data foreseen to be collected and how to connect with on-ground testing results and model simulations.
Highly Eccentric Orbits (HEO; eccentricities above 0.8 and perigee above the drag regime) are favourable for scientific and observations missions since the satellite is outside of the Earth’s radiation belt for most of the orbital period, avoiding noise and radiation effects which can interfere with Earth and Universe observations. Over the years, more than 30 missions have been operating in this type of orbits (i.e. Chandra, MMS, THEMIS amongst the US missions; ISO, XMM, Cluster, Integral are the ESA ones), and future missions in this type of orbits are planned such as the Plasma Observatory one.
For this particular class of orbits, the re-entry velocity (> 10 km/s) is higher than from a circular orbit, which translates in the radiative heat exchange as the dominant heat transfer mechanism. The strong bow shock wave generated in front of the re-entering body and the consequent temperature rise results in a non-equilibrium flow condition, leading to molecular dissociation. The complexity of this phenomenology implies chemical reactions which are expected to belong to observable spectra.
However, due to the low occurrences, HEO (or super-orbital) re-entries are usually studied in the context of the re-entry of meteorites rather than artificial objects and the gap of knowledge related to the phenomenology for this type of re-entries is still large.
The Cluster II mission is a constellation of 4 identical cylindrical satellites flying in a tetrahedron configuration on HEO. The re-entry of the first of these satellites is expected in early September. Due to the dominant third body perturbation, the re-entry trajectory is very stable and predictable to a large extent, providing the perfect target for an airborne re-entry campaign and a unique opportunity for a repeatable experiment.
The work presented highlights the knowledge gaps that Cluster-II re-entry observations aim to fill in terms of understanding the thermal response of satellite components and predicting breakup events.
The historic re-entry of a Delta-II upper stage in 1997 has been re-assessed using a simplified model in DRAMA v3.1 and SAMj. The analysis considers five scenarios covering both object-oriented and component-centric models of the spacecraft, as well as the impact of improvements in material representation since the case was originally executed in 2012.
The trajectories predicted by DRAMA and SAMj when operating in 3dof mode were found to be similar. The debris fields generated by both tools were significantly short of the along track distance seen in reality, which resulted in objects being found in Texas at approximately 30 degrees latitude. The results when using a component-centric model were slightly better than those generated by an object-oriented equivalent, with impacts at 56-59 degrees latitude, rather than 63-68 degrees latitude. Unfortunately, the simplicity of the spacecraft model meant that most components were not marginal from a demisability perspective, and therefore, the impact of material model improvements made over the last decade on surviving fragments was not significant.
Removing the 3dof tumbling assumption, and simulating the re-entry from a nose first attitude in 6dof using SAMj moves the debris field significantly further south to 34 degrees latitude. This occurs because the naive assumptions made about the distribution of mass within the spacecraft leads to a nose heavy model that is stable at this nominal attitude. However, it does consolidate the suggestion that early in the re-entry the spacecraft was consistently flying in a low drag orientation.
The benefit of evaluating the impact of uncertainties associated with re-entry events has been demonstrated within the activity through the execution of 2000 simulation Monte Carlo assessments for both a 3dof and 6dof case using ESA's PADRE framework. This manages the execution of re-entry Monte Carlos in DRAMA or SAMj and consolidates the results in terms of the aggregate risk, number of fragments, and their composition, impacting location, mass and energy characteristics. These analyses were seen to increase the range of predicted impact locations significantly, to 40-75 degrees latitude in the case of the 3dof evaluation and 15-65 degrees latitude when simulating with six degrees of freedom.
The latest data from 2022, shows a record number of satellites were launched and re-entered Earth’s atmosphere compared to previous years. Current projections that account for the growing prevalence of satellite constellations and launch rideshares, suggest that these numbers will continue to rise. This increase in atmospheric re-entry events shows the growing adherence to space debris mitigation measures, that mandate the removal of satellites from over-populated orbits at the end of their operational lives. However, as more satellites opt for disposal via atmospheric re-entry, the precise simulation of re-entry becomes increasingly important to assess the trajectory and associated risk of any surviving fragments.
Simulating re-entry scenarios is a complex process, involving the modelling of atmospheric, aerothermodynamic, structural and flight dynamics amidst shock wave interactions and high-temperature, non-equilibrium flows. Accurately modelling scenarios involving multiple interacting fragments also poses a distinct challenge to re-entry simulations. Current re-entry tools typically neglect the effects of fragment interactions, including the possible physical collisions and flow feature interactions. Neglecting these features may alter the profile of the fragment cloud post-breakup, subsequently influencing the trajectories and ground impact footprint predictions of re-entry simulations.
A re-entry tools ability to accurately simulate re-entry is dependent on the fidelity level of its models, particularly in the case of the fragmentation, aerodynamics and aerothermodynamics. Most re-entry tools rely on low-fidelity expressions for aerothermodynamics and aerodynamics, as well as pre-defined criteria for fragmentation, prioritizing computational efficiency over accuracy. In comparison, high-fidelity methods capable of resolving the complexities of re-entry are often considered computationally impractical when analysing full re-entry trajectories or in design for demise processes.
The re-entry analysis tool TITAN (TransatmospherIc flighT simulAtioN) addresses the aforementioned challenges of re-entry modelling through its collision model to simulate the physical interactions of fragments and its multi-fidelity capabilities for aerodynamics, aerothermodynamics and structural dynamics. This approach enables the use of high-fidelity models at critical phases of the re-entry trajectory to minimize the error of predictions. The collision model in TITAN allows for elastic collisions that accounts for the impact of object contact in the respective trajectory and risk analysis.
In this study, the destructive atmospheric re-entry of the Delta-II second stage is simulated with TITAN, with the collision model and multi-fidelity capabilities incorporated. The impact of including these models will be analysed with respect to the resulting prediction of the trajectory and ground footprint of the surviving Delta-II second stage debris. The results of the study will be assessed with respect to the Delta-2 rocket debris recovery locations to quantify the impact of the respective modelling methods on the accuracy of the predictions.
Over the last ten years, the prediction of space debris survivability during their re-entry and the associated prospective ground risk have received an increased interest in the scientific community due to complex multi-physics modelling requirements and crucial industrial applications. Nevertheless, important uncertainties in re-entry physics remain due partially to the lack of data and knowledge to calibrate such tools. Then the interest of developing more complex codes with coupled physics and achieving sensitivity analysis compared to real flight data makes perfect sense.
In this context and in the framework of the French Space Operation Act (LOS), to predict the debris survivability of a space vehicle and its associated fragments during their atmospheric re-entry, and assess the prospective risk on the ground, CNES in collaboration with R.Tech develops its spacecraft-oriented tool named PAMPERO. PAMPERO is a multidisciplinary tool that models a spacecraft's complete atmospheric reentry, including the fragmentation and ablation processes, along a six DOF trajectory.
Within the framework of the improvement/validation of PAMPERO and the in-depth understanding of the re-entry process, data from experiments is essential, regarding all the associated uncertainties.
This presentation aims to challenge the code PAMPERO on the Delta-II flight rebuilding using sensitivity analysis to deduce the most probable re-entry scenario.
SCARAB Numerical rebuilding Delta II Second Stage re-entry - ATD3 Test Campaign
As the number of small satellites in Low Earth Orbit grows, understanding their re-entry demise becomes increasingly crucial for ensuring the sustainability of space utilization. This includes risks associated with ground impacts from incompletely demised components and the release of aerosols and gases into the upper atmosphere. The SOURCE PWK project, funded by the German Aerospace Center (DLR) and carried out by the University of Stuttgart’s Institute of Space Systems (IRS), aims to investigate the demise behaviour of CubeSats. The investigation employs both numerical demise simulations of the CubeSat mission SOURCE (Stuttgart Operated University Research CubeSat for Evaluation and Education) and plasma wind tunnel (PWT) experiments on its critical components.
The project employs a two-pronged approach, combining numerical simulations of CubeSat missions like SOURCE (Stuttgart Operated University Research CubeSat for Evaluation and Education) with plasma wind tunnel (PWT) experiments on critical components. SOURCE, a 3U+ CubeSat developed by IRS in collaboration with the University of Stuttgart's small satellite student society (KSat e.V.), is expected for launch not earlier than 2025. Beyond educational and technological demonstration objectives, its mission includes gathering in-situ measurements during early re-entry phases above approximately 130 km altitude. These data can be used to refine existing numerical models and demisability analysis tools for spacecraft.
Numerical simulations using the ESA-code SCARAB, developed by HTG GmbH, preceded the PWT experiments, aiding in identifying critical components such as titanium threaded rods, antennas, printed circuit boards (PCBs), carbon fiber reinforced plastic (CFRP) elements, cameras, magnetorquers, and batteries. PWT experiments expose these components to high enthalpy air plasma flows simulating conditions at the stagnation point during re-entry. Trajectory points relevant for the demise process were selected from simulations, ranging from the early re-entry phase at approximately 93 km altitude to the peak heating phase around 80 km altitude.
The process of component demise in the PWT is monitored using a video camera (Sony Alpha 6400), an infrared thermal camera (FLIR A6751 SLS), a linear pyrometer (KE Technologie GmbH LP3), and thermocouples at specific points of interest. Optical emission spectroscopy (OES) is also performed in the stagnation point region using the visual and near-infrared range (OceanOptics HR4 VIS-NIR). The named instruments and measurement techniques can be combined to document the demise processes in a time-resolved manner. This allows for the correlation of the emissions of gaseous particles and droplets to the heating process of exposed components and materials.
This work presents the visual observation of the demise process, along with spatially and time-resolved temperature measurements and OES data for the tested components. The results are compared to the SCARAB re-entry simulations to identify significant deviations where demise models can potentially be improved. Finally, conclusions are drawn regarding the expected demise during SOURCE’s actual re-entry, specifically, and CubeSats in general.
The use of pyrotechnic formulations to assist spacecraft demise during re-entry has been investigated in the frame of the SPADEXO ESA-TRP project. The approach selected during this activity consisted in the direct integration of thermite powder in the structural voids of robust components, which could be passively ignited by the aerothermal heat experienced by the spacecraft during the re-entry process. The extra enthalpy released by the reaction could be exploited to induce or assist the demise, hence lowering the casualty risk on ground associated to uncontrolled re-entries. This concept, hereby named Thermite-for-Demise (T4D) was already presented in some papers [1-2] or preliminary studies [3], but this strategy was investigated for the first time in an extensive manner only during the SPADEXO project.
The activities of the project culminated in the experimental campaign performed at the L2K arc-heated hypersonic wind tunnel, located at DLR’s premises in Cologne. Steel mock-ups of particularly robust components were filled with thermite and exposed to the high-enthalpy hypersonic flow, to verify the thermite passive ignition and its effects on the samples. A dedicated extension of the SCARAB re-entry software was realized for the prediction and successive rebuilding of the results. After the experimental campaign, the samples were then analysed in detail at Politecnico di Milano, to characterize the combustion products and highlight the properties of the remaining slag.
The presentation will focus on the results obtained on the mock-ups of Solar Array Drive Mechanisms (SADM). These cylindrical samples were characterized by the possibility of embedding a thermocouple inside the lateral wall, whose data proved to be particularly valuable in quantifying the effectiveness of the thermite enthalpy release to the surrounding structure. The passive ignition verified in all the tested cases and the effectiveness of the heat transfer confirm the validity of the T4D concept, that is now looked with interest also for future commercial applications [4].
During the atmospheric entry, the decommissioned satellite is strongly affected by aerodynamic heating due to dissipation of a huge amount of kinetic energy into thermal energy. In these conditions, the vehicle usually breaks-up into several parts which are, in turn, degraded by the high enthalpy flow. However, Carbon Fiber Reinforced Polymers (CFRP) materials behave similarly to thermal protection ablators and, hence, show a strong resistance when exposed to high enthalpy flows. For instance, objects made with this type of material like Composite Overwrapped Pressure Vessels (COPV) often survive the re-entry. This component is made of a metallic liner wrapped in CFRP. During the atmospheric entry, the COPV will be exposed to a high enthalpy flow, which will progressively pyrolyze the resin of the CFRP, then erode by thermo-chemical phenomena the remaining carbonaceous residue together with the carbon fibers and finally melt the liner if the heat load is sufficient. To assess the performance of these tanks, two experimental test campaigns were conducted. One campaign investigated tank performance in a subsonic regime, while the other focused on the supersonic regime.
High-fidelity models have been developed in recent years to predict the response of light porous ablative thermal protection materials such as PICA and there is a growing interest in exploring the possible extension of those models to predict the demisability of space debris composite materials. The objective of this work was to perform an experimental and numerical campaign reproducing the demise of COPVs. Miniaturized COPV samples have been designed and manufactured to avoid delamination of the fiber layers. Those have been exposed to relevant atmospheric entry conditions in the Plasmatron facility at the von Karman Institute [1]. The subsonically tested samples were then simulated using the high-order discontinuous Galerkin code Argo. This employs a unified method, solving the surrounding flow and material response in the same domain of computation. It uses volume averaging theory to describe flow through the porous material and to capture with accuracy the gas-surface interaction. .
This methodology has shown its advantages to predict the response of low-density porous materials but its extension to very dense composite materials such as CFRP presents several numerical challenges that we will discuss. Experimental data and numerical results will be presented and compared. The interpretation of the results will aid in laying the foundation for a technological use of these procedures in the design phase of components susceptible to re-enter our atmosphere.
Reference:
1) J.ElRassi, B. Helber, P. Schrooyen, A . Turchi, P. Jorge, T. Magin, L. Walpot, Plasma testing of miniaturized composite Overwrapped pressure vessels in reentry conditions. 10th EUCASS Aerospace Europe Conference 2023. DOI:10.13009/EUCASS2023-957
An end-of-life scenario for the demise of a LEO satellite might start with deceleration from drag in the LEO environment, followed by heating, ablation, and breakup as the satellite descends into the dense atmosphere. Some key physical and chemical processes would be gas-surface energy transfer, ablation reactions on high-temperature surfaces, and pyrolysis of polymeric materials. Molecular beam methods can provide useful understanding of these processes. Molecular beam-surface scattering of O atoms and N2 molecules on representative satellite materials may be used to measure energy transfers and scattering angles, which are foundational to drag simulations. Molecular beam-surface scattering may also be used to investigate the oxidation reactions that lead to ablation on high-temperature surfaces. In another method, a pulsed hypersonic molecular beam with a high peak flux may be used to create a repeatable shock layer above a heated test article at 2-3 pulses/s, and recession measurements can be made. Such an experiment simulates the rarefied shock layer that forms above a re-entering satellite, allowing ablation phenomena to be studied as a function of material temp. A third method uses mass spectrometry to obtain the molar and mass yields of gaseous products as a function of temperature as a polymer or polymer-composite material is heated rapidly. With such data, finite-rate reaction pathways can be identified by observing the differences in temperature-dependent product yields at different heating rates. Summaries of all three experimental methods will be presented, with examples of gas-surface scattering dynamics of O atoms on representative satellite surfaces, fundamental studies of gas-surface reactions of O atoms on hot carbon surfaces, materials ablation phenomena of carbon subjected to repeated shock layers containing O atoms, and thermal decomposition of phenolic resin.
Operating satellites in very low earth orbit (VLEO), approximately between 100-300km altitude, has numerous advantages. One advantage is close proximity to earth, which can enable accurate sensing with significantly lower SWaP+C (size, weight, and power + cost). Another advantage is that VLEO is effectively a “self-cleaning-orbit” where unpowered spacecraft will quickly re-enter the atmosphere and burn up; thereby reducing or avoiding the space debris problem. However, operating in VLEO comes with significant challenges compared to conventional LEO satellites. Major challenges include atmospheric drag in the rarefied (and highly variable) upper atmosphere and material degradation due to atomic oxygen impacts at orbital velocity, as atomic oxygen becomes the most dominant species above 100km altitude. Future VLEO spacecraft will require drag-reduction strategies such as maintaining a ram-facing orientation that minimizes cross-sectional area and advanced materials/coatings that are both O atom resistant and low-drag. For these reasons, the scattering dynamics of high energy atoms and molecules on various materials becomes important to characterize.
In this presentation, a new gas-surface scattering model, that is able to reproduce a wide range of molecular beam surface scattering data on various satellite materials, is summarized. This scattering model is implemented within a three-dimensional direct simulation Monte Carlo (DSMC) code and simulations are performed to understand how an accurate scattering model influences the drag coefficient for various satellite-relevant geometries and the resulting influence on expected satellite lifetime in VLEO. Furthermore, the simulations predict local forces on all satellite surfaces and therefore moments about the center of mass. Combinations of surface orientation and local scattering dynamics can lead to either ‘aerodynamic’ stability or instability, which is important when trying to maintain a ram-facing orientation to minimize drag. Finally, the DSMC simulations are able to capture O atom buildup within open cavities and some exemplary results will be presented to quantify potential O atom damage on internal satellite components.
The aerodynamic forces and aerothermal loads experienced by spacecraft components during a fragmenting reentry likely influence their trajectories, demisability, and, as a consequence, the resulting ground casualty probability. Due to the complexity of this phenomenon, Design for Demise tools often simplify the dispersion of the fragments and the interactions between the components, considering their trajectories as independent (i.e. neglecting any mutual interaction) as soon as a structural limit triggers the fragmentation event. Studying the interaction of proximal bodies and clusters is required to develop improved separation models that can advance demisability predictions.
The present study (HiSST-2024-0142) investigates the aerodynamic separation of compact fragment clusters. Fragmentation scenarios assuming cluster compositions of equal-size spheres and cubes were analyzed. The aim of the investigation is twofold. First, to extend earlier sets of measurements towards higher Mach numbers in order to determine to which extent this parameter influences the flow separation velocities. Second, to evaluate to which extent the shape of the fragment can influence the dynamics of the separation process (spherical geometries have mostly been used to date, but might be poorly representative of a real fragmentation event). The experimental analysis was conducted in the VKI Longshot hypersonic wind tunnel at Mach 12 flow conditions. A dual-camera free-flight testing methodology was employed to observe separation scenarios and track the motion of the fragments in six degrees of freedom. The test articles are initially confined in a sabot, which separates into two pieces and exposes the models to the freestream upon the arrival of the flow. Two tests were performed on 11 spheres, two tests on 11 cubes, and one on 36 spheres.
Experiments on 11 sphere clusters yielded a reasonable agreement regarding the mean terminal velocities with the experimental observations of Whalen et al. 2021 (https://doi.org/10.1007/s00348-021-03157-z) and the correlation proposed by Park et al. 2020 (https://doi.org/10.1016/j.asr.2019.10.009). Increasing the sphere population to 36 yielded greater object spread. During both the 11 and 36 sphere tests, the formation of sub-clusters was observed, i.e., objects gathered behind a leading body. Similar observations were reported by Whalen et al. 2021. Tests on clusters of 11 cubes demonstrated significantly larger object dispersal, higher lateral terminal velocities, and larger maxima compared to the 11 sphere cases. These results indicate that the object shape contributes to the dispersal and should be accounted for in demisability predictions. The mean value and the distribution of the terminal velocities imply a reasonable consistency for the repeated experiments; however, such fragment separation is expected to be complex. Therefore, in order to draw clear conclusions and to represent adequately the separation dynamics on a macroscopic scale, a high number of repeat tests are required, and the problem must be assessed on a statistical basis.
The PRODUCERS project, funded by ESA and led by Fluid Gravity Engineering, Ltd., aimed to experimentally investigate and computationally predict radiative markers associated with the destructive re-entry of spacecraft at the end of their life. The purpose of the research activity is to improve the analysis of spectroscopic and visual data obtained through dedicated remote observation campaigns of such events. The primary reference case under consideration are the Cluster-II satellites. The experimental aspect of this activity involves a thorough investigation of spectral signatures related to the demise of materials and components in the Plasma Wind Tunnel (PWT) facilities at the Institute of Space Systems (IRS) at the University of Stuttgart.
Two key phases of the experiment campaign are summarised. The first phase focused on determining the spectrographic signatures of individual materials demising under aerothermodynamic heating conditions relevant to the scheduled Cluster-II re-entries. A selection of five different aerospace material samples, including CFRP, aluminum, titanium, and stainless steel alloys was subjected to destructive testing in the PWT with an emphasis on emulating enthalpies and pressures of key trajectory points around the stagnation area of the probe, with the aim of producing and capturing emissions from highly excited liquid and gaseous outflows, i.e. droplets. The corresponding temporally resolved emission signatures were measured using two optical emission spectrometers and correlated with visual and emissivity-corrected thermographic temperature measurements of the front surface. The results imply that material emissions, specifically those from metallic alloys can provide the means to identify both the composition and the ongoing state of demise of the observed materials, however it must be noted that the energy accumulation in the observed metallic droplets appeared to insufficiently emulate wake flow conditions for demising spacecraft, to the effect that emission lines associated with the primary element of a given alloy was rarely identified in the reported experiments.
The second phase of testing focuses on similarly assessing typical emission signatures of key spacecraft components and structures extracted from recovered as debris from the original Cluster spacecraft debris. The composition of such test articles is heterogeneous, resulting in richer emission signatures during demise as compared to homogeneous material tests. The tests demonstrate the challenge of distinguishing between different polymers based on their emission spectra, as most yield similar carbonaceous and organic material emissions. However, flare-ups recorded in the spectra can be noted as being indicative of certain macroscopic failure events. While the testing further implies that the spectral signature of small scale equipment may in principle permit an identification of sub-components e.g. in the case of electronics equipment, it is important to note that this is likely impractical during remote observations due to length scale considerations.
The experimental campaign created a comprehensive database of spectroscopic markers for various materials, surface coatings, functional and structural components that are relevant for the re-entry of Cluster-II and other spacecraft. The impact of PWT testing conditions and the measurement setup designs on the prospect of recreating spectrographic markers representative of observable flight conditions, i.e. wake flows, are discussed.
Setting up a clean scenario for space management, where space debris are effectively removed from orbit following a design-to-demise strategy presents a formidable challenge. This involves employing design tools commonly developed and used for Thermal Protection System (TPS) of reentry vehicles. However, designing for space debris demise introduces a higher level of complexity, requiring engineers to adhere to much more stringent criteria and account for additional physical phenomena.
Several codes have been developed to initiate the design process and establish preliminary parameters for space debris demise. Nevertheless, these codes necessitate further consolidation and validation to ensure their efficiency and accuracy in real-world scenarios. To accomplish this, experimental test cases must be meticulously addressed to simulate the intricate conditions of reentry and assess the performance of satellite designs under such circumstances.
Central to this exploration is the identification and utilization of ground facilities capable of providing relevant testing conditions for evaluating spacecraft demisability. The selection of appropriate test facilities and test conditions are pivotal in accurately determining the effectiveness of engineering solutions designed to facilitate the controlled demise of space debris.
The present study aims to identify the capacities and the limits of aerospace ground testing facilities to realize destructive testing of typical satellite equipment. It will be the occasion to present how much high-enthalpy and plasma wind tunnels could offer full scale testing in full reentry conditions. Following this first analysis ground testing strategies for space debris re-entry duplication will be assessed for destructive testing of spacecraft equipment. The conclusion will open the discussion on the dedicated validation framework needed to properly consider experimental data and D4D tools.
TA6V (90% Ti, 6% Al and 4% V) is a titanium alloy currently used in spacecraft construction, due to its high mechanical strength and high corrosion resistance. In order to limit the damage caused by space debris, the oxidation phase during earth atmospheric re-entry has to be studied. As TA6V is used in a wide range of applications, many articles were already published on its resistance to oxidation under standard air [1-3]. However, there are only few articles on its oxidation under air plasma [4, 5]. Yet during earth atmospheric re-entry, oxygen is dissociated and diffuses more easily into the material. As titanium is highly reactive to oxygen, it can significantly modify the development of the oxide layer.
In this work, TA6V samples were oxidized by air plasma generated in the SOUPLIN inductively coupled plasma torch (60 kW, 1.8 MHz) at the CORIA laboratory. Air pressure and anode voltage are the two parameters that modify plasma intensity. Thus, one sample is oxidized by a low-intensity plasma (1410 Pa, 5 kV) and another by a high-intensity plasma (4530 Pa, 6 kV). The oxidation time is set at 1 min.
The oxidized samples were then analyzed using various complementary characterization techniques such as X-ray diffraction, Raman spectroscopy and scanning electron microscopy coupled with energy dispersive X-ray spectroscopy. The chemical nature and the distribution of the oxides developed were studied on the surface and in the depths of the material.
To assess the effectiveness of tools designed for hypersonic material testing and space debris analysis, it is essential to replicate specific environmental conditions such as pressure, temperature, and shear stress in ground-based testing facilities. As part of the latest developments in space debris material research, the Plasmatron facility at the von Karman Institute now provides a unique setup enabling off-stagnation testing within a supersonic framework [1]. In this study, ground tests were carried out on woven CFRP (Carbon Fiber Reinforced Polymer) and titanium samples using the VKI-Plasmatron to emulate the conditions encountered by these materials during reentry into Earth's atmosphere as space debris. Three samples of each material were tested in a supersonic plasma flow, either parallel to the flow or with a slight angle of attack (AoA). Various intrusive and non-intrusive instrumentation techniques were employed during testing to monitor the thermal response and oxidation growth of the materials and to visualize the temperature distribution across the surface of the samples.
The CFRP samples was made to represent rocket fuel tanks, so called wrapped COPV (Composite Overwrapped Pressure Vessel). The interaction between strong forces ripping off the carbon fibers and their simultaneous oxidation is crucial to the fuel tank's demise during re-entry. The surface temperature monitoring shows a rapid initial temperature rise until the pyrolysis of the resin starts. While the pyrolysis is ongoing the temperature stays steady, and slowly rises as the pyrolysis ends if the heat flux is sufficient. A clear increase in the recession speed was observed with an increase in angle of attack and higher flow enthalpy. The material reached a steady surface temperature under ablation, and the effect of shear was clearly pulling the fibres off the sample. The titanium samples showed a complex oxidizing behaviour, varying for different heat fluxes and shear conditions. In the low shear test case without angle of attack, a protective layer of white TiO2 formed. This brittle later remained on the sample. With higher shear, increasing AoA, this layer was ripped off, and a complex active oxidation occurred. The oxidation could be tracked in time with the cameras. First a blue oxide layers travels from the sample front all the way to the back of the sample. As the samples heats up further, this oxidation layer degrades and is sheared off, and the active oxidation starts. More in depth material research is needed to characterize these surface oxides.
This study aims at describing the physio-chemical alterations in standard layered materials, analogous to those utilized in spacecraft construction, under the influence of reactive atomic oxygen (ATOX). ATOX conditions similar to those found in Low Earth Orbit (LEO) were simulated to assess, quantitatively and qualitatively, the impact of ATOX exposure [1] in mass loss, structural and chemical transformations.
Pristine samples were also compared as reference in order to understand the effect of ATOX. Additionally, post-exposure changes in thermo-optical properties, including the materials' solar absorption and reflectivity, were examined.
The research findings illuminated the primary degradation patterns of the layered materials. The study provided crucial insights into the potential particle sizes that could be released during atmospheric re-entry. Moreover, the observations established a correlation between the mechanical degradation of the materials and changes in their thermo-optical properties.
With the recent awareness of the space sector on the fragile near-Earth space region and the forecast of the booming number of satellited objects, various mitigation approaches are currently evaluated and start to be implemented to limit the impact of space activities and achieve a safe and sustainable space environment.
The re-entry of spacecraft into the Earth’s atmosphere can contain fragments which are able to survive the loads and heat experienced during re-entry into the atmosphere.
Numerous studies and projects have been performed in the past in order to gain better understanding of re-entry processes and to assure that more spacecraft fragments demise during re-entry manoeuvres.
From recent simulation activities under ESA contracts, the list of potential S/C critical objects identified Star-trackers optics and electronics, as equipment of interest. Therefore the objective of this Demise for Demise activity is to enhance the knowledge on the demise of optical and electronic equipment of satellite platforms in order to establish validated re-entry models.
Plasma wind tunnel test campaign with dynamic set-up has been performed in order to better assess fragmentation and behaviour of the aforementioned components.
The activity will aid in improving the modelling and simulation of spacecraft component demise behaviour, and further industries understanding and capabilities in more sustainable spacecraft design.
The Kentucky Re-entry Universal Payload System (KRUPS) provides a quick-turnaround, low-cost plat- form to conduct atmospheric entry experiments. KRUPS is designed to test multiple types of thermal protection systems (TPS) and scientific instrumentation. Five KRUPS capsules were sent to the International Space Station (ISS) via the NG-20 Cygnus resupply vehicle. After the completion of the resupply mission, the Cygnus vehicle de-orbited with the capsules inside. Cygnus then broke up into the atmosphere in order to burn up stored trash. These five capsules constitute the second Kentucky Re- entry Payload Experiment (KREPE-2) mission, each with a different heat shield TPS material. Following on the success of the first KREPE mission, the second generation of capsule design added updated avionics, extended battery life, and more scientific instrumentation. Added instrumentation included an updated flight computer, 5 port flushed air data sensing (FADS) pressure port system, GPS receiver, pre-calibrated IMU, and a spectrometer. In addition to this added instrumentation, the KREPE-2 capsules can transmit back 5 times more scientific data than the first generation KREPE-1 capsules via the Iridium satellite network. This data will help with the reconstruction of the atmospheric entry environment and validation of computational fluid dynamics (CFD) and material response (MR) models developed at the University of Kentucky. With the completion of KREPE-2 on the horizon, the design of the seven capsules of KREPE-3 just started, with one of them potentially used to study demise.
Spent upper stages of launch vehicles are a very large and dangerous class of space debris since they are prone to spontaneous explosions producing many small objects. This makes old upper stages one of the primary targets for future active space debris removal missions. If the initial attitude motion of the target stage is known, e.g., from in-situ measurements exploiting CubeSats, it is possible to attach a deorbiting device to the stage in order to ensure its safe deorbiting and reentry.
This study proposes concepts of passive devices that can change the character of the attitude motion of the target following two different deorbiting and reentry scenarios, the choice between which depends on the specific mission’s parameters. The first scenario is aimed at minimizing the debris footprint size. In this case, it is necessary to ensure that the stage does not tumble during reentry, i.e. that its longitudinal axis is closely aligned with the velocity vector of the center of mass. To achieve this, the deorbiting device needs to have fins oriented in such a way as to make the stage spin. The second scenario is intended to ensure that the fragments produced by the eventual breakup of the stage do not reach the Earth’s surface. In this situation, the breakup should occur at high altitudes, thus, unlike the previous case, spinning is undesirable, but tumbling is necessary since it provokes large aerodynamic and thermal loads on the stage, forcing its early breakup. To despin the stage in the initial phase of the mission, the deorbiting device can use a yo-yo mechanism, and tumbling can be achieved if the device is equipped with a fin on one side, which is perpendicular to the stage’s longitudinal axis.
The applicability of the above-mentioned scenarios and concepts of deorbiting devices has been demonstrated by numerical simulations of the attitude motion of the Arian 4 upper stage during reentry.