ATD3 Workshop 2026

17 Jun 2026, 07:30 → 19 Jun 2026, 23:00 Europe/Amsterdam

Accueil (Toulouse - Cité de l'Espace)

Aurora Saracini, Cristina De Persis (ESA), Dominique Pirotais, Julien Annaloro (CNES), Pierre-Henri Maire, Stephane Galera

 

AeroThermoDynamics and Design for Demise (ATD3) Workshop

ATD3 workshop is a long-standing European forum which aims at fostering technical and scientific discussions, collecting and disseminating knowledge, proposing new topics and activities, and contributing to planning efforts, including roadmap definition and coordination.

The main objectives of the ATD3 workshop are to:

• Establish a framework for the verification, validation, and comparison of numerical methods used in space object reentry simulation tools.
• Disseminate recent scientific and technical results within the ATD3 community, which includes academic participants (in particular PhD students) as well as experts from industry.
• Coordinate and discuss future activities, with the goal of gathering expert perspectives from both academia and industry on forthcoming developments in the fields of aerothermodynamics and design for demise

The workshop addresses a broad range of topics related to space object re-entry aerothermodynamics and design for demise, including:

• Improvement, verification and validation of numerical methods for simulating the destructive re-entry of space vehicles, including verification of design for demise solutions;
• Modelling of ablation and material response under re-entry conditions;
• Fragmentation and break-up modelling of re-entering space objects;
• Experimental capabilities, testing facilities and measurement techniques supporting re-entry analysis;
• Characterisation of materials relevant to atmospheric re-entry;
• Investigation of drag-driven de-orbiting solutions.

ATD3 2026: preliminary information

The next ATD3 workshop will be held from June 17^th to June 19^th 2026 in Toulouse (France). The event will be hosted at the Cité de l’Espace (Space City).

The format is a sequential list of presentations of about 20 minutes each, with 5 minutes brief Q&A session per presentation.

We are proud to have HyFAR-ARA as our official sponsor, whose support makes this event possible. Thank you for partnering with us to create an exceptional ATD3 experience.

Preliminary agenda

 

Important deadlines

 

17 May 2026: Extended abstract submission deadline 

25 May 2026: Abstract selection announcement

27 May 2026: Preliminary program online

1 June 2026: Registration deadline

17-19 June 2026: Workshop

Wednesday 17 June

Registration

ATD3 Workshop Opening: Welcome and Opening Speeches

Numerical modelling and validation of destructive re-entry

1 Tree-Based Surrogate Modelling of Aerothermochemistry for Space Debris Reentry

Accurate aerothermodynamic modelling of space debris reentry is critical both for ground
casualty risk assessment and for quantifying ablation-driven mass deposition in the
upper atmosphere. Operational tools, such as ESA’s DRAMA software suite,
rely on engineering correlations with simplified chemical assumptions, while high-fidelity
CFD, though resolving some of these limitations, remains prohibitive for the probabilistic
trajectory ensembles required in Design-for-Demise workflows. Moreover, part of the
chemistry occurring during reentry remains difficult to capture even at high fidelity. This
motivates the development of fast, physically grounded surrogate models that can serve
as a foundation for progressively more complex coupled aerothermochemical problems.
A 0D reference solver is constructed to generate the training database. Post-shock
conditions are obtained from the Rankine–Hugoniot relations solved iteratively via
Newton–Raphson, with thermochemical equilibrium computed through Gibbs free-
energy minimisation using the Mutation++ thermochemical library. Wall boundary
conditions, for a carbon-based spherical object, are handled through a Surface Mass
Balance (SMB) formulation accounting for carbon injection via sublimation (C3
formation), oxidation, and oxygen catalysis. This yields mass blowing rate, heat flux
partition, and near-wall species mass fractions. The input space spans altitude 30–80 km,
Mach 20–28, and body radius 0.1–1.0 m, covering the continuum reentry corridor for low-
Earth-orbit debris.
Three tree-based ensemble regressors (Decision Tree, Random Forest, and XGBoost) are
benchmarked in a unified multi-output architecture predicting eleven quantities
simultaneously: total heat flux, mass blowing rate, electron number density, and nine
near-wall species mass fractions. Hyperparameter optimisation is performed via Optuna,
a Bayesian framework based on the Tree-structured Parzen Estimator, with 3-fold cross-
validation and automated median pruning to ensure that observed performance
differences are attributable to algorithmic structure rather than tuning disparity.
XGBoost achieves R²scaled = 0.999976 on the test set, outperforming Random Forest
(0.999935) and Decision Tree (0.999720). Optuna-tuned XGBoost reduces MAEscaled by
a factor of three relative to a default baseline, with heat flux identified as the hardest
output through a composite difficulty index combining R², MAE and RMSE across all
targets.
These results establish tree-based ensembles as a practical and competitive algorithmic
family for aerothermochemical surrogate modelling, currently underexplored relative to
ANNs and kriging in the debris reentry literature. The unified multi-output architecture,
providing heat flux, blowing rate, and ablation product composition simultaneously, is
directly amenable to embedding in higher-fidelity reentry codes.
This work is conducted within the ESA OSIP PhD fellowship framework, which aims to
predict element-specific atmospheric mass deposition profiles (such as Al, Cu, or C3 ).
The present surrogate constitutes the first building block of that chain. Validation of the
modelling approach against plasma wind tunnel data (VKI Plasmatron) and in-flight
optical observations is planned as next steps, targeting a first operationally viable
environmental footprint assessment within a reentry modelling chain.

Speaker: Jeanne Longlune (Thermo and Fluid Dynamics (FLOW), Faculty of Engineering, Vrije Universiteit Brussel & Aerothermomechanics department, Ecole Polytechnique de Bruxelles, Université libre de Bruxelles)

2 On Debris Aerodynamic Stability and its Prediction

In 2008, France adopted the French Space Operation Act (LOS), which regulates the space activities of French operators. Among many topics covered, the LOS act also addresses the reentry of spacecraft and their debris, by imposing a maximum acceptable probability of a fragment to survive and pose a risk to populations.

As the number of satellites in orbit continues to grow, the importance of accurately predicting debris behavior is also increasing. Thanks to all the extensive work performed by the scientific community, significant progress has been made in recent years in predicting the survivability of space debris during reentry. However, prediction of some aerothermodynamical phenomena remains challenging.

One of these phenomena is the aerodynamic stability, which has a high impact on the behavior of debris during their reentry and especially on their attitude. This, in turn, directly influences the heat flux distribution and ablation. Debris rotating at high rates has a different heat flux distribution compared to a fully stabilized body, which receives most of the heat on the side exposed to the freestream.

The aerodynamic stability is mainly characterized by pitching, yaw and roll moment coefficients and by corresponding moment damping coefficients. While the moment coefficients are well predicted with modern theories for capsules and other vehicles, some geometries like cylinders or rings show a noticeable disagreement when using simple models like modified Newton one. Moreover, the prediction of moment damping coefficients presents a significant challenge, due to the lack of related data and due to the dynamic aspect of damping.

The aim of this presentation is to share a methodology for analyzing the stability of reentering bodies and to highlight the limitations of Newtonian theories in predicting aerodynamic stability by comparing analytical solutions, CFD and available wind-tunnel data. The method was verified using a space capsule, and then the methodology was applied to a cylinder and a ring to analyze their stability. Finally, different functions of pressure distribution were tested to determine whether the prediction of moment and damping coefficients could be improved when using analytical models.

Speaker: Alexey KLIMKO (CNES)

On Debris Aerodynamic Stability and its Prediction - Abstract.pdf

3 BLAZE project: Improving Re‑Entry Predictions Through Cross‑Validated Modelling

Assessing spacecraft demisability during atmospheric re‑entry is essential for ensuring compliance with Space Debris Mitigation Requirements. Today, ESA’s DRAMA‑SARA toolchain is the de‑facto standard for such analyses. However, its empirical modelling assumptions and legacy conservatism increasingly drive major design decisions, often leading to unnecessarily constrained propulsion architectures, mass budgets, and disposal strategies. Recent studies at OHB indicate that DRAMA is being used far beyond its intended scope and tends to systematically overestimate survivability and casualty risk.

In our FAR 2025 contribution, we introduced OHB’s first end‑to‑end, physics‑based re‑entry modelling framework using DSMC solvers (dsmcFoam+ and SPARTA), augmented by detailed atmospheric (NRLMSISE‑00) and reactive chemistry (QK models). These simulations, applied to both capsules and full satellite geometries, revealed convective heat fluxes up to 2 to 4 times higher than DRAMA predictions at key breakup altitudes, clearly demonstrating the tool’s underlying conservatism.
Building on these results, the internal BLAZE project (Breakup & Layered Analysis of Re‑Entry) extends the comparison across the full ecosystem of re‑entry solvers (DRAMA, DEBRISK, SCARAB, PAMPERO, SACRAB) and introduces a structured cross-validation methodology jointly developed with RTech.

This contribution presents the combined results of the FAR 2025 work and the ongoing BLAZE campaign, highlighting the implications for spacecraft design and European industrial competitiveness. We show that high fidelity DSMC/CFD simulations provide realistic and physically consistent heat loads across the transitional flow regime, allowing to compare high fidelity tools with lower fidelity tools. The tool-to-tool comparison performed in collaboration with R.Tech provides more insight in potential differences between different medium and lower fidelity tools available in Europe.

The findings strongly support the need for an open discussion on DRAMA validation, conservatism, and requirements formulation within ESA’s Clean Space framework. Ultimately, our work aims to enable more balanced, evidence‑based demisability assessments that ensure safety without imposing unnecessary design penalties on future European missions.

Speaker: Bayrem Zitouni (OHB)

4 Risk Requirements Verification Methods and Models for Spacecraft Re-entry: DIVE Update, Explosion and Ablation Modelling

Ensuring safe and sustainable re-entry of space systems has become an increasingly important topic within European space debris mitigation efforts, especially with the evolution of Design for Demise (D4D) techniques increasingly demanding physically representative aerothermodynamic, fragmentation, and material modelling.

Existing re-entry tools, such as ESA’s DRAMA suite, provide robust capabilities for trajectory propagation and casualty risk estimation, but still rely on simplified or heritage assumptions for fragmentation mechanisms, explosion triggering and material ablation under re-entry aerothermodynamic conditions.

In recent years, several ESA-funded activities have significantly improved our understanding of demisable materials, hardware design solution, and re-entry modelling approaches. However, these advances have not yet been fully incorporated into the current version of the Design for Demise Verification Guidelines (DIVE).

The AERIS project addresses this gap by both updating the DIVE guidelines and improving the underlying modelling capabilities used in re-entry analyses. The activity is structured around three main pillars: (i) consolidation and update of DIVE based on recent experimental and modelling studies, (ii) development of a probabilistic explosion model for re-entry scenarios, and (iii) implementation of an ablation model capable of capturing material degradation and by-product release.

The first activity focuses on providing a structured update of the DIVE guidelines, building on the most relevant results from ESA studies carried out between 2020 and 2025. This update is based on a systematic review and consolidation of available studies, covering areas such as advanced materials, demisable hardware concepts, and approaches to testing and validation. One of the key challenges lies in the consistency of the inputs, as different studies may provide overlapping or sometimes conflicting recommendations. To address this, a structured approach is adopted to assess the relevance and reliability of each input and to ensure that all updates to DIVE remain traceable and technically justified.

In parallel, the project supports the evolution of re-entry modelling capabilities through the development of two complementary models meant to be used in re-entry analysis tools (component-oriented, e.g., DRAMA, or more accurate alternatives). A probabilistic explosion model is introduced to represent both the likelihood and consequence of in-flight break-up events happening during the atmospheric re-entry. In addition, an ablation model is developed to provide improved estimation of material degradation during re-entry.

Overall, AERIS supports more consistent D4D verification by aligning recent developments with the guidelines already present in DIVE. This is complemented by the integration of new explosion and ablation models to improve physical representativeness.

Speakers: Ms Anca-Maria Stan (Indra Space), Andreea Sabau (Indra Space), Federico Bariselli (von Karman Institute for Fluid Dynamics), Marco Fossati (University of Strathclyde), Valentin Ledermann (R.Tech Engineering)

Risk Requirements Verification Methods and Models for Spacecraft Re-entry - DIVE Update, Explosion and Ablation Modelling Abstract.pdf

15:50 Coffee & Networking Break

ATD3 Workshop 2026 Group Photo

Numerical modelling and validation of destructive re-entry

5 Reachable Set Propagation for Space Debris Re-Entry

As the Low Earth Orbit (LEO) satellite population increases, methods for computation of space debris re-entry footprints are an increasingly important part of assessing the impact of a spacecraft on humanity. Unfortunately, methods for forward and backward propagation of uncertainty and state estimation in uncontrolled re-entry struggle with a great deal of the common complicating factors on uncertainty propagation. Re-entry dynamics exhibit a strong dependence on attitude, so are thus high-dimensional, and are strongly non-linear and therefore also feature significant non-Gaussianity. The simulators applied to such problems also carry significant computational expense. Hence the feasible methods in this domain must make sacrifices in terms of physical or statistical accuracy, according to the well-known bias-variance tradeoff. The common preferred choice is to be statistically conservative through use of Monte Carlo campaigns.

Here an alternate approach is proposed that is computationally efficient and maximally conservative in terms of problem uncertainty due to the sacrifice of statistical information. Based upon the concept of reachable sets from control theory, a methodology is introduced that propagates the bounds of a volume of reachable space in the state space of the problem. By requirement of an aerodynamic model where aerodynamic coefficient ratios can be predicted based upon attitude, i.e. panel codes, and assumption that all uncertainties can be bounded by intervals the number of necessary samples for an absolute worst case debris footprint can be drastically reduced.

By selecting angle-of-attack, angle-of-sideslip pairs on the wind frame 2-sphere, attitude states can be found that result in aerodynamic forces that maximise reachability. Combining the state parameters with the uncertainty space and applying specific assumptions on re-entry models, a bounding box on the reachable set can be propagated.

This method is applied to a re-entry case of an upper stage with uncertainty in terms of initial state, atmospheric conditions and modelling assumptions and compared to a conventional Monte Carlo propagation. Finally, consideration is given to inverse problems and backward set propagation.

Speaker: Tommy Williamson (University of Strathclyde Aerospace Centre for Excellence)

6 High-Fidelity Coupled Flow-Structure Simulation of the early Reentry Phase of the Satellite Eu:CROPIS

Over the past decades, near-Earth space is becoming increasingly congested thus increasing the risk of inter-collision between satellites and generating thousands of space debris. To keep the near-earth orbit usable, it is necessary to deorbit the satellites at the end of their lives. However, during atmospheric reentry, parts of the satellite might not burn up completely and thus pose a non-negligible risk to ground safety. Typical risk assessment tools for satellites reentry are based on simplified geometrical and physical assumptions, limiting their capabilities to accurately capture local aerothermal effects.

Especially, inner components of satellite are shielded from the hypersonic flow during the early phase of the reentry, delaying their burn up. The use of demisable joints, as developed in the TEMIS-DEBRIS (TEchnologies for MItigation of Space DEBRIS) project, to artificially open up earlier the structure, appears promising to enhance the global demisability of satellites. Indeed, exposing the internal components to the outside hypersonic flow earlier would enable an earlier temperature rise and thus enhance the burn up.

To get a better understanding of the early reentry phase (ranging from 120km to 100km), this study applies high-fidelity numerical methods to investigate the reentry behavior of the Eu:CROPIS satellite.
DSMC (Direct Simulation Monte Carlo) simulations are applied to the satellite Eu:CROPIS reentry and coupled with a structural FEM solver to investigate the aerothermal heating of the satellite during early reentry. A non-reactive three species mixture (O₂, N₂, O) flow is simulated assuming isothermal walls and diffuse reflection with full thermal accommodation. An aerothermal database, which serves as the input to the structural solver, is then built from the surface heat flux on 11 trajectory points ranging from 120km to 100km altitude. The coupling between flow and structure is achieved using Conffass framework (Coupled Numerical Fluid, Flightmechanics And Structure Simulations).
High fidelity FEM structural simulation, accounting for both outer and inner radiation, are performed, simulating the transient thermal response of the satellite along the reentry trajectory. Initial temperature of the whole satellite is set to -23°C at the starting altitude 120km.

Peak surface heat fluxes from the DSMC simulations reach up to 83 kW/m² at 100km altitude on the windward faces, driving the rapid heating of the outer structure while inner components remain shielded. The coupled flow-structure simulation shows that the inner components of the satellite are indeed shielded during the early stage of the reentry, thus delaying greatly the burn up of them. While the outer structure of the satellite quickly reaches melting temperature of the aluminum at 660°C after 1840s at 115km altitude, the inner components remain comparatively cold like the inner payload which temperature does not excide 80°C.

These findings highlight the importance of accounting for structural shielding in reentry risk assessments. They also enable the setting up of a demise strategy by placing the demisable joints, which are thermosensible, at strategic locations to maximize the break-up altitude.

Speaker: Lukas Lemaitre (DLR)

Broad Scope Abstracts

7 The Future Launchers Preparatory Programme’s Aerothermodynamics and Demisability activities

ESA’s Future Launchers Preparatory Programme (FLPP) is dedicated to advancing Europe's space transportation capabilities: to, from and in space. It supports the development and testing of new technologies and system concepts that will shape the next generations of European space transportation systems. From reusable launchers and in-space transportation to advanced propulsion, the future launchers preparatory programme prepares the future of space transportation. This talk will introduce FLPP, highlight some of our activities related to Aerothermodynamics and Demisability, and explain the opportunities to work with FLPP in the future.

Some examples of FLPP’s projects on demisable materials for use in space transportation vehicles include:

• Developing a thermoplastic composite material which will ablate during re-entry. The project consortium includes a launch vehicle component manufacturer to provide the use case to be considered in the project, materials characterisation experts who will test the material in an ablation chamber which simulates re-entry conditions, and a startup specialising in Life Cycle Assessment to quantify the environmental performance.
• Developing a novel composite made from enhanced wood, with similar mechanical properties to aluminium, which will reduce the metallic particles being ejected in the upper atmosphere during re-entry.

FLPP also worked with DLR’s Supersonic and Hypersonic Technologies Department on a project to mature and validate Wind Tunnel Testing and Computational Fluid Dynamics for retro-propulsion, with the goal of preparing tools needed for the design of future European reusable launchers. It resulted in a number of new capabilities for aerodynamics and aerothermodynamics, for example being able to show the difference in heating of the base of the vehicle during a re-entry burn between firing a single engine and multiple engines, which is strongly affected by the plume-plume interaction. We are now running a follow-on project with the same team, which started by inviting the European Space Transportation primes to propose study cases which will directly address their current challenges and at the same time allow the consortium to mature and validate Wind Tunnel and CFD techniques. Although the cases proposed by industry mostly address aerodynamics topics, one of the selected cases will study aerodynamic heating of a multi-engine microlauncher during the ascent phase.

FLPP runs a number of different types of calls for proposals, including THRUST! for high thrust engines and BEST! for reusable boosters. Most relevantly to the current topic, the Future Innovation and Research in Space Transportation! (FIRST!) calls help technology providers to develop disruptive technologies which have commercial application in Space Transportation. Each FIRST! starts with a call for ideas through the OSIP platform, followed by a pitch day where the people who proposed the most interesting ideas can meet with Space Transportation primes and potential investors. An Invitation To Tender is then issued, resulting in between 6 and 10 contracts for one-year development activities targeting TRL 4 to 5.

With the support for Space Transportation shown at CMIN25, FLPP looks forward to continuing to prepare for the future of Space Transportation for the benefit of Europe.

Speaker: Dr David Riley (European Space Agency - Future Launchers Preparatory Programme (STS-FF))

8 Assessment and Mitigation of Ground Casualty Risks from Space Activities

The increasing number of space operations and orbital objects raises growing concerns regarding the risk of ground casualties associated with atmospheric re-entries and other space activities. Within the French regulatory framework, safety assessments and risk mitigation measures play a central role in the authorization and oversight of space operations.

This presentation will provide an overview of how a national entity (France) adopts the necessary and up-to-date regulatory measures to ensure the safety of space operations under its authority and thus meets its commitments under international treaties. More precisely, it will deal with approaches used to assess and mitigate ground casualty risks from space activities, with a particular focus on the French operational experience and regulatory considerations. It will discuss the interaction between technical analyses, compliance verification, re-entry monitoring activities, and evolving practices supporting sustainable and safe space operations.

Speaker: Mr Grégoire LAUR (CNES)

Thursday 18 June

08:30 Coffee & Networking

Ablation and material response

9 Eliminating B′ Tables: Time-Accurate Explicit Coupling in PATO for Hypersonic Solvers

Tightly coupled simulations between ablative materials and hypersonic flow solvers are essential for reliable demise-oriented design. The traditional use of B′ tables to couple material response and flow behavior was an effective strategy to address the computational limitations of the 1960s, but it remains limited in both accuracy and flexibility. This work proposes a generic coupling approach that leverages modern computational capabilities to provide a simpler and more accurate alternative. An explicit coupling boundary condition has been implemented in PATO resolving at each material time-step heat and mass transfer and chemistry at the interface—and through for porous materials. As a proof of concept, Mutation++ is used for non-equilibrium surface interactions, coupled with the hypersonic solver Eilmer to resolve the flow. This approach enables the simultaneous consideration of ablation, mass transfer, pyrolysis gas release, and blowing effects, which strongly influence the heat flux and species composition at the fluid–solid interface. Tested on representative hypersonic cases, the methodology represents a step toward a multiphysics framework capable of capturing material degradation, cracking, and fragmentation during atmospheric reentry. PATO is released as open-source software and is designed to be readily coupled with hypersonic flow solvers.

Speaker: Mr Gabriel Merlaud (Université de Bordeaux)

resume_ATD3_eng_jl.odt

10 Multi-scale stochastic modeling of porous ablators

Charring ablators protect spacecraft by coupling low thermal conductivity and endothermic pyrolysis with porous outgassing to produce transpiration cooling. This work establishes an end-to-end multiscale modeling framework for these TPS materials. At the pore scale, we performed detailed DSMC simulations of high-temperature rarefied gas flow through reconstructed fibrous preform geometries. The DSMC results predicted the permeability of these fiber networks in continuum/transitional regimes, matching CFD/theory and laboratory data. Crucially, coupled DSMC simulations with outward-blowing pyrolysis gas and O-atom diffusion showed that the outgassing strongly curtails oxygen penetration (to only ~0.2–0.4 mm depth), significantly reducing net oxidation and surface recession.

At the continuum scale, a finite-volume material-response solver (KATS) was deployed that captures full three-dimensional, anisotropic behavior of porous ablators. As part of this development, we introduced a 3D transient pyrolysis-gas transport model coupled to an orthotropic thermal-conductivity model for the charred composite . This fully coupled solver integrates conductive heat transfer, internal pore-gas convection, and surface pyrolysis/oxidation kinetics in one framework. The studies demonstrated that including internal gas flow and directional conductivity significantly alters predictions of surface temperature and recession relative to simpler 1D or isotropic models. In practice, the macroscale simulations use closure parameters (effective permeability, conductivities, etc.) obtained from the pore-scale DSMC analyses, ensuring consistency across scales.

This strategy tightly couples modeling and experiment across scales. Micro-CT imaging and flow-tube tests supply pore-scale geometry and material properties used in the models, while microscale simulations yield the constitutive relations needed by the continuum solver. For example, it was recently demonstrated that through these multiscale simulations, we were able to match the experimental permeability of fragile TPS preforms, directly informing the simulation inputs. In summary, the multiscale approach blends pore-resolved DSMC, novel material characterization, and 3D continuum CFD into a predictive framework. The integrated results capture how porous microstructure, pyrolysis outflow, and coupled ablation physics combine to determine heat-shield performance.

Speaker: Savio Poovathingal (University of Kentucky)

11 Quantification of stratospheric solid aerosols injection by global rocket launch and spacecraft re-entry

Over the last twenty years, space activities have expanded rapidly, marked by a threefold increase in launch rates and a thirtyfold increase in the number of satellites deployed (Taupin et al., 2025). In 2025, the mass of anthropogenic material re-entering the atmosphere was estimated to represent between 10% and 32% of the natural cosmic influx, with aluminum contributions alone exceeding cosmic source influx (Carrillo-Sánchez et al., 2020; Ferreira et al., 2025). This increase is largely driven by objects orbiting below 600 km altitude, where orbital lifetimes rarely exceed a decade.
When these objects ablate during atmospheric reentry, most of them inject gases and solid aerosols that are deposited primarily in the mesosphere and stratosphere. If they accumulate in large quantities, these aerosols can influence atmospheric composition, ozone chemistry, and radiative processes across various spatial and temporal scales (Ferreira et al., 2024; Ross et al., 2014). It is therefore essential to accurately determine past and current injection rates to assess their environmental impacts.
In this study, we present an alternative framework for estimating the mass released through ablation during atmospheric re-entry of spacecrafts. Previous approaches have either relied on theoretical average ablation coefficient (Schulz et al., 2021) or targeted individual chemical species (Ferreira et al., 2025). Here, we generate a range of ablation scenarios using the DEBRISK software developed by CNES, applied to simplified representations of real satellite models, lower and upper stages with varying masses, cross-section and orbital characteristics. These simulations are coupled with object-level information from DISCOSweb to reconstruct the cumulative mass injected into the stratosphere over the 1981–2020 period.
Furthermore, we created a new dataset describing specific launcher properties and propellant usage in order to quantify emissions from rocket launches, including black carbon and alumina particles. Altitude dependent injection profiles are calculated using representative flight trajectories and propellant burn rates for solid and liquid-fueled launch vehicles.
Finally, the resulting estimates are briefly compared with trends derived from the NASA Cosmic Dust Catalogs to evaluate the consistency between modeled inputs and in-situ observations of solid aerosol populations in the stratosphere.

References:

Carrillo-Sánchez, J. D., Gómez-Martín, J. C., Bones, D. L., Nesvorný, D., Pokorný, P., Benna, M., ... Plane, J. M. (2020). Cosmic dust fluxes in the atmospheres of Earth, Mars, and Venus. Icarus, 335, 113395.

Ferreira, J. P., Huang, Z., Nomura, K. I., Wang, J. (2024). Potential ozone depletion from satellite demise during atmospheric reentry in the era of mega-constellations. Geophysical Research Letters, 51(11), e2024GL109280.

Ferreira, J. P., Wang, J. (2025). Determining mass fluxes of space debris upon demise in the atmosphere. Acta Astronautica.
Ross, M. N., Sheaffer, P. M. (2014). Radiative forcing caused by rocket engine emissions. Earth’s Future, 2(4), 177-196.

Schulz, L., Glassmeier, K. H. (2021). On the anthropogenic and natural injection of matter into Earth’s atmosphere. Advances in Space Research, 67(3), 1002-1025.

Taupin, Q., Lasue, J., Määttänen, A., Zolensky, M. (2025). Constraining the origins of terrestrial stratospheric solid aerosols over the 1981-2020 period (No. EPSC-DPS2025-1899).

Speaker: Quentin Taupin (CNES - IRAP - LATMOS)

12 Investigations of High-Temperature Oxygen-Carbon Ablation by Molecular Beam-Surface Scattering

An end-of-life scenario for the demise of a LEO satellite might start with deceleration from drag, 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. We have used molecular beam methods to provide insight into all these processes, yet this presentation will focus on the findings of many molecular beam-surface scattering experiments to gain an understanding of the high temperature (1000 – 2000 K) atomic-oxygen-induced ablation of model carbon materials. Earlier molecular beam data were used as the basis of an air-carbon ablation model (generally referred to as the “ACA model”), but refinement of this model was needed to make it more applicable to a variety of carbon types and environments. Thus, new molecular beam experiments have been conducted to expand the range of carbon materials studied, and the new data have been provided to inform an update of the ACA model in the group of Prof. Tom Schwartzentruber at the University of Minnesota.
The new experiments were performed with pulsed molecular beams of O atoms. The reactive scattering dynamics of O on various carbon surfaces – vitreous carbon (revisited), HOPG, isostatically-molded graphite (IMG), and a 3D carbon-carbon composite (C/C) – suggest that the oxidation mechanisms on all sp2 types of carbon are similar but that the morphology of the surface region strongly influences the relative importance of the individual mechanisms. The relative probabilities of the observed gas-surface interactions, both reactive and non-reactive, were quantified as a function of surface temperature. Carbon surfaces that have lower density or more porosity tend to exhibit higher reactivity. In addition to reacting with carbon to produce CO2 (minor product) and CO (major product), unreacted O atoms may scatter through thermal and non-thermal processes, and, with sufficiently high incident O-atom flux, O-O recombination on the surface to produce O2 may occur with an efficiency that is somewhat lower than that to produce CO. Furthermore, O atoms may be retained through adsorption on the surface or absorption into the bulk, leading to slow production of CO on a timescale of tens of microseconds to hundreds of milliseconds. The slow processes through which CO is formed and released from the surface may account for as much as half the CO that is produced at lower temperatures, whereas the slow production of CO becomes negligible at higher temperatures and all the CO is released within a few microseconds of O-atom impingement on the surface. Relative to vitreous carbon, IMG and C/C show broader temperature ranges over which slow thermal release of CO remains important. The molecular beam scattering results indicate the importance of microstructure in the temperature-dependent ablation rate of carbon by atomic oxygen. The results further show that slow CO production, on a timescale up to hundreds of milliseconds, should be considered in experiments and models where the finite rate of oxygen-carbon ablation may be important.

Speaker: Timothy Minton (University of Colorado Boulder)

ATD3 2026_Minton.pdf

13 From Molecular Beam Studies to Aerodynamics: Development of a Gas-Surface Scattering Model for LEO/VLEO

Spacecraft end-of-life mitigation strategies are increasingly important to tackle the space debris problem. Among these, drag-driven de-orbiting solutions such as deployable drag sails rely on increasing aerodynamic drag to accelerate orbital decay. Alternatively, operating satellites in very low Earth orbit (VLEO) offer a passive de-orbiting approach, where the denser atmosphere rapidly leads unpowered spacecraft to re-entry at end of life. However, the dense atmosphere produces significant drag during normal satellite operations, requiring novel surface materials/coatings which promote gas-surface scattering consistent with low platform drag. In both drag sail de-orbiting and VLEO operations, predicting spacecraft orbital decay requires accurate aerodynamic models.

In these rarefied environments, aerodynamic coefficients can be calculated using direct simulation Monte Carlo or similar numerical methods. The aerodynamic coefficients calculated with these methods fundamentally depend on the gas-surface scattering models used to describe the scattering dynamics of atmospheric particles impinging on the spacecraft. However, commonly used gas-surface scattering models are not sufficiently detailed to capture material-specific surface properties, leading to uncertainties in the calculation of aerodynamic coefficients.

This presentation provides an overview of a new gas-surface scattering model, detailing how it can be parametrized using molecular beam experiments, and subsequently used to estimate aerodynamic coefficients. The gas-surface scattering model uses a stochastic representation of the surface and includes surface corrugation as an input parameter. The model parameters are fitted using molecular beam-surface scattering experiments, which measured the scattering dynamics of hyperthermal atomic beams of argon and oxygen on satellite surfaces with different roughnesses. We observe that the fitted roughness parameters correlated positively with the surfaces topographical parameters measured by atomic force microscopy. While the fitted corrugation parameters are not identical to the AFM-measured roughness, the positive correlation between them supports the model’s physical basis. These scattering measurements are then linked to macroscopic aerodynamic forces: the surface-specific fitted parameters are used in the gas-surface scattering model to estimate the aerodynamic coefficients of flat plates, coated with the tested surfaces, at different angles of attack. The resulting values illustrate how different materials and surface morphologies alter aerodynamic behavior under orbital conditions.

Speaker: Pedro Jorge (University of Colorado Boulder)

Abstract_ATD3.pdf

10:30 Coffee & Networking Break

Ablation and material response

14 Assessment of Particle Formation During Re-entry of Spacecraft

There is a substantial lack of knowledge about the environmental effects to the upper atmosphere by the continuously increasing number of satellites that are brought to re-entry after operation. The main constituent of satellite structures is aluminum. Aluminum is injected into the Earth’s atmosphere as a rather new element, as it is only a minor constituent in micrometeorites [1]. However, the impact of this new trace element remains scarcely investigated so far.

Current research assumes the immediate oxidation of molten or evaporated aluminum during re-entry to form aluminum monoxide (AlO). Then, upon cooling, the formation of alumina (Al2O3) particles [2,3], or aluminum hydroxides (Al(OH)x) is discussed in literature [4]. However only few experimental data are available of the processes, especially on the forming solid particles. In our group, we are trying to experimentally evaporate aluminum and collect the resulting solid particles after they eventually cooled down again.

These experimental simulations are performed in the plasma wind tunnels at the Institute of Space Systems (IRS) at the University of Stuttgart. In past experiments we observed the evaporation of aluminum and the formation of AlO. By adding aluminum powder in the plasma flow, we acquired spectral signatures of gaseous Al and AlO, similar to those measured during airborne observations of re-entries [5]. We suspect that the aerothermal environment plays a major role in the particle formation processes. For example, a simple steady evaporation of particles seems not to cover the aerodynamic effects of resident times of particles in the hot core flow. A very first analysis of presumably molten or evaporated particles impacting on copper plates in different locations is used to study the formation processes. The impacting aluminum formed a thin layer coating the copper. Preliminary Powder X-Ray Diffraction and Scanning Electron Microscope analyses by the Department of Chemistry of the University of British Columbia in Vancouver [6,7] show that the layers consist of Al2O3 particles and pure metallic aluminum with a thin top layer of Al2O3.

In the presentation we will explain the newly developed experimental setup to study the processes after the demise of re-entering satellites. We assume several factors influence the process from the spacecrafts’ demise towards the formation of solid particles. These experiments are of high interest to gain an understanding of the increasing number of satellites re-entering the earth’s atmosphere.

[1] Schulz and Glassmeier, Advances in Space Research, 2026.
[2] Park and Layland, Acta Astronautica, 2021.
[3] Maloney et al., JGR Atmospheres, 2025.
[4] Plane et al., JGR Space Physics, 2021.
[5] Loehle et al., Meteoritics and Planetary Science, 2021.
[6] https://www.chem.ubc.ca/x-ray-crystallography
[7] https://emlab.mtrl.ubc.ca/equipment/

Speaker: Dominik Kuenstler (High Enthalpy Flow Diagnostics Group (HEFDiG), Institute of Space Systems (IRS), University of Stuttgart)

15 Finite-Rate Ablation and Oxide-Layer Models for Reentry Spacecraft

Reentry spacecraft experience extreme heating and surface chemistry that heatshield materials must withstand or that can lead to the demise of spacecraft materials and release of product species into the atmosphere. A common assumption made in such analysis is equilibrium surface chemistry. While this may be accurate for carbon recession rates under certain conditions, it does not accurately predict product species. Furthermore, equilibrium surface chemistry is likely not accurate for oxide layer growth and volatilization, which involve slower chemical reactions compared to carbon oxidation. For these reasons, finite-rate ablation and oxide layer models are needed that are accurate over a wide range of temperatures and pressures experienced by reentry spacecraft materials.

Recently, molecular beam experimental data was used to create the finite-rate air-carbon ablation (ACA) model [1]. By explicitly including surface coverage effects, the model captures the non-Arrhenius trend for CO production versus increasing temperature and also introduces pressure dependence. New comparisons between CFD simulations using the ACA model and recent carbon ablation experiments will be presented [2]. Next, a reformulation of the ACA model will be presented that can predict ablation of various forms of carbon by two controlling parameters; the number of available reactive sites and the ratio of strongly-to-weakly bound oxygen on the surface. The new ACA model formulation is compared to new molecular beam data for HOPG, vitreous carbon, and graphite materials.

Modeling oxide layer formation and volatilization is important for reusable hypersonic materials and also for predicting the demise spacecraft. Some thermal protection systems form silica-based oxides and most metals will form complicated oxides when exposed to oxygen at high temperature. Previous modeling has mainly used equilibrium chemistry approaches, however, the time history of the oxide layer may be important as its growth and/or volatilization occurs on timescales comparable to trajectory variations. We therefore implemented the finite-rate model of Fertig et al. [3] into the US3D CFD code and verified against prior test case results. New results will be presented that demonstrate finite-rate oxidation layer formation in a high-enthalpy reactive air flow. Both finite-rate models (carbon ablation and oxide layer formation) and the coupling to CFD tools may be useful for predicting the demise behavior of carbon-based materials and oxide-layer-forming materials upon reentry.

[1] Prata, K.S., Schwartzentruber, T.E. and Minton, T.K., “Air–Carbon Ablation Model for Hypersonic Flight from Molecular-Beam Data”, AIAA Journal, 60(2), pp.627-640, 2022.
[2] McClernan, P.G., Schroeder, O.M., Fagnani, A., Knutson, A.L., and Schwartzentruber, T.E., "Finite-Rate Modeling of Air–Carbon Ablation in a Plasma Wind Tunnel, " JTHT, Vol. 40, No. 2 (2026), pp. 323-336 doi: doi/abs/10.2514/1.T7237
[3] Fertig, M., Herdrich, G., and Auweter-Kurtz, M., “SiC Oxidation and Catalysis Modelling for Re Entry Heating Predictions,” European Space Agency, (Special Publication) ESA SP, Vol. 659, Jan. 2009

Speaker: Tom Schwartzentruber (University of Minnesota)

Schwartz_ATD3_2026.pdf

16 Modeling Spacecraft Demise Using the KATS Multi-Physics Framework

Accurate prediction of spacecraft demise during atmospheric re-entry requires tightly coupled modeling of aerothermal loading, material response, and structural failure. This work presents the capability of the Kentucky Aerothermal and Thermal-response Solver framework (KATS) to address this challenge through an integrated, multi-physics approach. The methodology combines KATS-FD for flowfield reconstruction, KATS-MR for detailed material response, and KATS-SM for structural mechanics, enabling a unified simulation of degradation and breakup processes under hypersonic conditions.

KATS-FD provides the external aerothermal environment, capturing high-enthalpy flow effects and surface heat flux distributions along the trajectory. These loads are directly coupled to KATS-MR, which resolves in-depth material behavior including pyrolysis, oxidation, and ablation for a range of thermal protection and structural materials. The resulting thermochemical state and recession rates are then passed to KATS-SM, a structural module that incorporates a crack formation and propagation algorithm to model progressive weakening and fragmentation of the structure.

A key feature of the framework is its ability to predict the onset and evolution of structural failure as a function of both thermal degradation and mechanical loading. The crack formation model in KATS-SM enables simulation of fracture initiation driven by thermal gradients, internal pressure, and material property evolution, providing a physics-based pathway from intact structure to fragment generation. This approach allows for more realistic prediction of breakup altitude, fragment size distribution, and release conditions compared to traditional uncoupled or empirically driven methods.

The integrated KATS framework therefore offers a powerful tool for design-for-demise studies, enabling improved assessment of spacecraft survivability and debris risk. By bridging flow physics, material response, and structural failure within a single modeling environment, the approach supports the development of more reliable predictive capabilities for controlled and uncontrolled re-entry scenarios.

Speaker: Alexandre Martin (University of Kentucky)

atd3_KATS.pdf

17 Developing a Pyrolysis Gas Thermal Blocking Model for Reentry Demise

In NASA’s Object Reentry Survival Analysis Tool (ORSAT), aerodynamic drag and aerothermal heating coefficients are computed for each of the free-molecular, continuum, and transitional flow regimes using analytical and semi-analytical methods. These heating coefficients were derived for typical metallic materials that melt and do not have a strong gas-phase contribution to the flow in the boundary layer. Modern satellites typically feature fiber-reinforced polymer (FRP) components, such as solar array booms, facesheets of sandwich panels, or overwraps for composite-overwrapped pressure vessels (COPV). These FRP materials do not behave the same as metals in the reentry environment, but instead will pyrolyze and develop significant volumes of gas into the boundary layer.

Accurately predicting the reentry demise of FRP components is critical to assessing the reentry casualty risk for modern spacecraft. Research in recent years has shown that this demisability can depend heavily on how the expulsion of gaseous pyrolysis products through the outer surface of the material affects the heat flux at the surface. The ODPO has been developing a reduced-order model of the effect of pyrolysis gas blowing on the heat flux based on correlations between a blowing factor and a non-dimensional heat flux to be incorporated in the upcoming version 7.3 of the Object Reentry Survivability Analysis Tool (ORSAT). This presentation discusses the progress of this development project and the challenges remaining for generalizing the model across families of FRP materials.

Speaker: Dr Benton Greene (NASA ODPO (JETS-II Contract))

ESA Reentry Workshop Abstract_final.pdf

18 Thermal and mechanical responses of a pyrolyzable heat shield subjected to ablation and deformation

During atmospheric hypersonic re-entry, the heat distribution within the thermal protection system (TPS) is dampened by the in-depth chemical degradation of materials - called pyrolysis -, and by a surface physico-chemical degradation - called ablation. The aim of this work is to enhance pyrolysis modeling by considering solid deformations in order to describe more accurately the solid geometry variations resulting from swelling and ablation. The further aim of this study is to ensure mass and energy conservation during the pyrolysis-thermal coupling of heat shield under deformations. First, an overview of macroscopic modeling of pyrolysis is done. Arrhenius laws are employed for the density variation prediction. Then, thermal expansion, swelling and shrinkage are investigated as a consequence of material degradation, in addition to ablation. This analysis explores a pyrolysis-thermal model preserving mass and energy conservation during deformation and a number of numerical resolution maintaining numerical conservation. Finally, the model and methods are validated on ablation and swelling test cases from the literature and then applied to in-house experimental cases. The simulation results are in reasonable agreement with reference data and experimental data. Including swelling provides a closer approximation of wall evolution during hypersonic re-entry simulation.

Speaker: Céline Baranger (CEA)

CNES aerothermodynamics Activities Overview

12:45 Lunch Break

Material characterization for re-entry applications

19 Thermochemical study of Ti-6Al-4V alloy for material characterisation during atmospheric re-entry

The rapid increase in space activities has led to a growing number of objects re-entering the Earth’s atmosphere, raising safety and sustainability concerns. Thus, reliable prediction of re-entry behaviour, including aerothermal loads, ablation, fragmentation, and survivability, depends strongly on accurate material characterisation. However, accurate data remain limited for materials exposed to high-temperature, reactive, and partially ionized environments, especially for metallic alloys.

For atmospheric re-entry, material characterisation requires a detailed understanding of thermodynamic behaviour, phase changes, oxidation mechanisms, and chemical composition evolution over a broad temperature range. Interaction with the atmosphere at several thousand kelvins leads to complex gaseous species and partially ionized plasma, strongly affecting heat transfer and ablation. Despite advances in thermochemical databases and modelling tools, data for multi-component metallic alloys remain sparse compared with classical thermal protection materials.

This work analyses the thermodynamic, transport, and compositional behaviour of metal–air mixtures to identify the main mechanisms governing material response during re-entry. Alloy Ti-6Al-4V is considered in the study. This material is widely used in aerospace applications, including structural components such as propellant tanks. It has high mechanical strength and good corrosion resistance. This makes its behaviour and survivability under re-entry conditions particularly relevant. Equilibrium thermochemical analysis is performed using the Mutation++ library. The metal-air interaction is analysed over various compositions, from metal-rich to air-dominated cases at atmospheric pressure and in the temperature range 500 – 6000 K.

The results reveal a transition from a Ti-dominated, metal-rich regime at low air fractions to chemically driven behaviour as air content increases. At 2500 – 4500 K, oxide formation becomes significant, mainly through TiO and TiO₂, while nitrides appear in smaller amounts. These species progressively dissociate at higher temperatures. The response is strongly non-linear, especially above 80% air, where small composition changes produce large thermodynamic effects. Equilibrium specific heat shows multiple peaks linked to oxidation, nitride formation, and dissociation, which sharpen and shift to lower temperatures with increasing air content. Dynamic viscosity varies smoothly, whereas thermal conductivity increases strongly at high temperatures due to reactive contributions. Above 5500 – 6000 K, all cases tend toward atomic species, indicating strongly dissociated regimes.

This work represents a first step toward building datasets for metallic and composite materials, supporting future data-driven and machine-learning approaches for re-entry modelling. Future work will focus on extending the analysis beyond equilibrium assumptions, incorporating gas-surface interactions, and expanding the database to a wider range of materials.

Speaker: Ella Barakhovskaia (1 Faculty of Engineering, Thermo and Fluid Dynamics (FLOW), Vrije Universiteit Brussel (VUB), Brussels, Belgium; 2 Brussels Institute for Thermal-fluid Systems and Clean Energy (BRITE), Vrije Universiteit Brussel (VUB) and Université Libre de Bruxelles (ULB), Brussels, Belgium)

20 Introduction of innovative material solutions for demisable propellant tanks during atmospheric reentry

Hydrazine tanks for Low Earth Orbit satellites are nowadays constituted of Ti-6Al-4V titanium alloys. The constitutive alloy is subjected to constraints of different order. First of all, it must support the inner pressure without breaking. Then, it should be demisable during atmospheric reentry. Indeed, at the end of a satellite life, an atmospheric reentry operation is mandatory to limit the total amount of space debris. The current propellant tanks are only partially demisable, which induce high casualty risks for humans and the environment.
New materials should thus be considered to fabricate the tank. Different materials families were considered in this work (Ti alloys, TiAl intermetallics, Al alloys). To evaluate their respective demisability, numerical simulations of atmospheric reentry on a critical trajectory were performed on test cases, using both DEBRISK and ARES softwares - respectively developed by CNES and ONERA. The results obtained with DEBRISK are averaged on the total mass of the tank, whereas the results obtained with ARES are local, and provides the maximum values locally reached by the tank. Consequently, simulations allow to determine some key materials parameters (melting temperature, heat of fusion, thermal conductivity and emissivity) to be adjusted for a better demisability. Moreover, previous researches have shown that developing frangible materials is a potential solution to improve the demisability of the tank.
An experimental approach was then developed to evaluate the potential of three material solutions to substitute the current Ti-6Al-4V alloy. First, Ti-6Al-4V alloy was fabricated using additive manufacturing (Electron Beam - Powder Bed Fusion), followed by a consolidation thermal treatment. The aim is to achieve an architecture made of alternated porous and dense areas. Frangibility of titanium alloys could also be obtained thanks to eutectic structures. Indeed, alloying titanium with β-eutectoids stabilizer elements like Ni, Fe, Si or Cu creates an eutectic microstructure whose melting temperature is significantly reduced compared to titanium alloys. Some eutectic alloys (Ti-Ni and Ti-Fe, with different alloying elements compositions) were fabricated by arc melting. Their characterization constitutes a necessary step to further work on architecturated titanium eutectic alloys, which could be processed via powder metallurgy routes. Finally, intermetallic TiAl alloys were considered as potential material solutions. Alloying TiAl with Nb and arc melting elaboration creates a solidification structure constituted of a refractory Nb rich squeleton, which could induce frangibility.
A material characterization campaign was then led on the fabricated samples. Mechanical behaviors were evaluated by 4-points bending tests. Thermal and thermodynamic properties (conductivity and emissivity at high temperature, heat capacity) were measured for each sample and compared with the results available in the literature. Finally, oxidation tests in laboratory air (for experimental conditions – time, temperature - representative of an atmospheric reentry) were performed. The different oxide layers have been identified as they are of primary importance for the thermal emissivity of metals. Simulations using experimental results as input for material properties are finally made for a numerical validation of the three material solutions which were considered.

Speaker: Benoit Fer (ONERA)

Abstract_ATD3.pdf

21 T4D - Mechanical and energetic characterization of DLP printed thermite

In the context of Design for Demise (D4D), there have been multiple attempts at incorporating energetic thermite powders inside hard-to-demise spacecraft components, to provide additional energy during re-entry. This philosophy is called Thermite for Demise (T4D).
Those attempts highlighted a multitude of problems and inefficiencies arising from the use of thermite in the form of loose powder.
In the EIC project THREAD, thermite consolidation strategies are being explored to solve this problem.
For this purpose, stoichiometric magnesium - silicon dioxide thermite was consolidated via Digital Light Processing (DLP) 3D printing with a solid loading of 60%. The composition was chosen for its overall favourable compatibility with the selected technology, based on powder granulometry, energy band gap, and lack of toxicity.
The complete production procedure has been fully detailed and reported.
The properties of the material were characterised in terms of ignition behaviour, burn rate, chemical kinetics, mechanical properties, friability, and Electrostatic Discharge (ESD) safety.
Mechanical properties were assessed by performing compression tests, obtaining results in terms of Young modulus and yield stress. Tumbling tests were employed to evaluate the behaviour of this material during handling and transport, a scenario where previous experiments conducted on pellets of simply pressed thermite showed remarkable fragility and a tendency to easily crumble.
Energetic properties were assessed by Differential Thermal Analysis (DTA) for the kinetic parameters, as well as for reaction onset temperatures at very low heating rates.
Igniting the samples on a resistive strip, heated by a controlled current source, allowed to recreate heating rates typical of re-entry. Temperature readings were performed using a pyrometer focused on the heating element, at the very base of the thermite sample.
The combustion reaction was captured via a high speed camera to determine the burn rate of the thermite under those conditions.
To guarantee the safety of people operating with this new material, ESD sensitivity testing was performed up to an energy of 0.05 J.
The material's mechanical properties were surprisingly almost unaffected by the high mass loading, when compared with samples of pure resin binder. Cylinders of the material also demonstrated great resilience to tumbling, losing little to no powder during the tests.
Precise controlled ignitions in the inert atmosphere of the DTA machine revealed an inhibited reactivity, when compared to the original loose powder, which is consistent with the presence of the binder. Ignitions on heated strips revealed a pulsed ignition behaviour, where the material alternated shedding its binder via pyrolysis and thermite ignitions in multiple steps. The burning rate was measured for each subsequent ignition event. The ignition temperatures recorded during said tests were consistently higher with respect to the ones seen during DTA analysis, underlying the difficulty of matching ignition data across those two diagnostics.
ESD testing revealed that both the loose powder and the consolidated one have no sensitivity to electrostatic discharges, an important feature for operational safety.
In conclusion, thermite consolidated via DLP 3D printing shows promising results in the creation of custom energetic components for possible future satellite demise applications.

Speaker: Carlo Zanardi (Politecnico di Milano)

15:30 Coffee & Networking Break

Experimental capabilities and ground testing

22 Demise Experiment on the Upcoming KREPE 3 atmospheric re-entry mission

Design-for-demise strategies require accurate understanding of the thermal, mechanical, and fluid environments experienced by spacecraft during uncontrolled atmospheric re-entry. However, in-flight data characterizing the internal conditions of a disintegrating host vehicle remain extremely limited. The KREPE-3 mission, part of the Kentucky Re-entry Universal Payload System (KRUPS) program, provides a unique opportunity to address this gap through a dedicated demise experiment embedded within a multi-capsule re-entry architecture.

The experiment consists of a compact instrumentation package integrating an inertial measurement unit (IMU), pressure sensors, and thermocouples, designed to operate from the initiation of the de-orbit maneuver through the period immediately preceding capsule ejection (KREM separation). This measurement window captures the transition from orbital conditions to the onset of aerothermal loading and structural degradation within the host vehicle. The objective is to quantify the internal environment experienced by payloads during early-stage breakup and to characterize the conditions governing their release.

Particular emphasis is placed on reconstructing the dynamics of capsule ejection. Current debris survivability and dispersion models rely on simplified or assumed initial conditions at release, introducing significant uncertainty in predicted casualty risk and footprint. By correlating kinematic data from the IMU with local pressure and temperature measurements, the experiment aims to establish realistic initial states for multiple embedded objects at the point of separation.

In parallel, the mission includes the KRACO vehicle, a subscale analogue of the ESA DRACO capsule, intended to support design-for-demise validation under flight conditions. KRACO is designed to de-risk DRACO through two primary objectives: (i) evaluation of spectral emission markers to enable reliable airborne optical tracking and identification during re-entry, and (ii) assessment of vehicle stability through the transonic regime, with particular attention to center-of-gravity placement and its impact on attitude behavior. These measurements will provide critical data for correlating ground-based observations with in-flight dynamics and for validating stability predictions in a regime where aerodynamic uncertainties remain significant.

The resulting dataset will support the validation and improvement of demise and debris propagation models, including those used in design-for-demise assessments. In addition, the measurements will provide insight into the timing, sequencing, and variability of payload release from a degrading structure under hypersonic conditions. These findings are expected to contribute to more reliable prediction of fragment behavior, improved spacecraft design practices for controlled demise, and enhanced compliance with emerging space safety guidelines.

The KREPE-3 demise experiment represents a step toward bridging the gap between ground-based testing, modeling assumptions, and true flight conditions for spacecraft breakup and debris generation.

Speaker: Alexandre Martin (University of Kentucky)

atd3_KREPE3.pdf

23 Preparing DRACO for Re‑entry: VKI’s Integrated Design, Testing, and Material Demise Assessment

DRACO (Destructive Re-entry Assessment Container Object) is an ESA Space Safety mission designed to obtain the first in-situ measurements of a spacecraft’s fragmentation and demise during atmospheric re-entry. A washing machine-sized satellite will be intentionally deorbited, carrying approximately 200 sensors and multiple cameras to monitor structural breakup, thermal loads, and material behaviour. All data will be transmitted to a 40 cm survival capsule engineered to withstand the spacecraft’s disintegration and continue broadcasting during its parachute-assisted descent. Scheduled for launch in 2027, DRACO will provide unprecedented real flight data to improve re-entry modelling and support ESA’s Design for Demise and Zero Debris objectives by revealing how satellites actually disintegrate in the upper atmosphere.

The von Karman Institute for Fluid Dynamics (VKI) has played a central role in DRACO from Phase A through the consolidation Phase B/C, contributing to capsule design, aerothermodynamic modelling, and the experimental validation of critical subsystems. VKI’s work includes the definition and testing of the Thermal Protection System (TPS), the reentry capsule, the development of the Thermal Insulation System (TIS) for the onboard computer, and the demise characterisation of representative spacecraft materials and their instrumentation. These activities collectively strengthened the mission’s technical maturity and provided the evidence base required for Phase C consolidation towards CDR.

During Phase B/C, VKI executed an extensive plasma wind tunnel campaign to qualify the TIS and Sample Materials (SM) under conditions representative of DRACO’s re-entry environment. The TIS was tested against the mission-defined requirements, and multiple configurations were evaluated, including Sigratherm MFA-coated panels, MFA FF/Dalfatherm layups, a fully instrumented TIS box with a camera window, and a three-dimensional “corner” configuration. Across all tests, the TIS maintained structural integrity, coating adhesion, and acceptable back face temperatures.

The SM campaign investigated the demise behaviour and spectral signatures of materials expected on the DRACO platform. Tested items included lanthanum hexaboride marker material, additively manufactured titanium, titanium with embedded spectral markers, and coated titanium. The tests revealed clear emission signatures (La, Na, Ti, and oxide species) and documented melting, oxidation, and fragmentation processes relevant for re-entry demisability modelling and optical identification strategies. Additional tests will be carried out on instrumented structural elements of the satellite platform, such as aluminium L-profiles and anodised sandwich panels, including thermocouples, strain gauges, and heat flux sensors.

Speaker: Dr Bernd Helber (von Karman Institute for Fluid Dynamics)

24 Space debris testing methodology in Plasma Wind tunnel for D4D tools validation

See document attached

Speaker: Olivier Chazot (VKI)

ATD3_SpaceDebrisDemise_PWT_June2026.pdf

25 KLARA: AN ARIANEGROUP-CNES INITIATIVE FOR MONITORING SATELLITE FRAGMENTATION.

More and more objects are present in space, but what happens to them when entering the atmosphere? With the exponentially increasing number of satellites, it is essential to minimize the risk of falling elements in habited areas. To best predict what happen, it is necessary to predict with the highest accuracy the behavior of objects during their atmospheric reentry and destruction. Models exist for this, but observation and measurement in real conditions are not yet available. To this end, within the leagacy of the development of capsules from the Space Case family[1] [2], developed for different experiments in flight and calibration of multiphysical models, a version named KLARA is being developed in cooperation with CNES. This effort directly supports the needs identified in the CNES roadmap in the frame of the validation of DEBRISK /PAMPERO certification tool

Speaker: gregory pinaud (ArianeGroup)

ATD3_2026_AGS_KLARA.pdf

26 Increasing Demise Through Design: Application of Shape Effects

There are a number of design-for-demise methodologies which can be employed to increase the demisability of spacecraft components, which reduces the ground risk from re-entry. One of the less used concepts is to increase the heating to a component. Active methods, such as the use of energetic materials, and passive methods such as the use of lattice structures have been attempted with varying degrees of success. Recent high enthalpy demise ground testing has shown that heating to objects is significantly affected by length scales and flow paths. This suggests that simple geometric features such as holes, grooves and steps have the potential to increase the heat flux received by a spacecraft component during re-entry, with minimal change to the component.

In order to test this hypothesis, a set of simple shape adaptations have been tested in the DLR H2K cold hypersonic wind tunnel. Four concepts were selected based on sparse literature data on the likely effectiveness; holes, facets, steps and grooves. The test involved the use of a PEEK model, with infra-red thermography used to measure the temperature change at all locations on the model. As PEEK is a well characterised material in terms of emissivity and conductivity, the heat flux can be inferred from this data. A one-dimensional semi-infinite wall assumption was used, which will have some errors at sharp corners, but the methodology provides a good estimate of the heating to the features tested.

The grooves, steps and facets concepts were tested by adapting a cylinder to include these features. Tests were performed with the cylinder axis normal to the incoming flow, and at 45 degrees to the incoming flow. All three concepts showed a significant increase in the peak heat flux, and both steps and grooves produced a significant increase in the overall heat flux. The facets were less effective, resulting in a similar integrated heating over the surface. Grooves were the most effective of these concepts, both in terms of the integrated heating, but also in terms of the effectiveness at essentially any angle of attack. This is due to the grooves continually providing new leading edges such that the boundary layer growth is limited. Preliminary indications are that heat flux increases of 50% could be possible, which is significant in terms of the demisability potential of materials such as steel.

Holes were tested on a sphere cap. Three different hole sizes were used, and significant heating is observed downstream of the hole. Clusters of holes have also been shown to be highly beneficial, with smaller gaps between the holes increasing the effectiveness. Closed holes were also tested, and showed some benefit, but significantly less than through holes.

Two of these concepts, holes and grooves, will be tested further to demonstrate their effectiveness in a plasma wind tunnel.

Speaker: Ian Holbrough (Belstead Research Limited)

Social Event: Dinner

Friday 19 June

08:30 Coffee & Networking

Fragmentation modelling

27 ESA-DSMCFED: Development Status and Recent Advances of the SMURFS Re-Entry Fragmentation Code

At the end of their operational lifespan, uncontrolled spacecraft re-enter Earth's atmosphere and undergo total or partial destruction through interaction with the surrounding flow. The growing population of debris in Earth's orbits has prompted space agencies to address the space debris issue by imposing increasingly stringent requirements over the years. The published ESA standard Space Debris Mitigation Requirements (ESSB-ST-U-007) recommends Design for Demise (D4D) as the preferred end-of-life mitigation strategy for limiting the risk of casualties on the ground. The D4D philosophy advocates that, early in the design and development stages of a new space mission, careful evaluations of spacecraft buses, payloads, and structural components be conducted to assess their potential to survive an uncontrolled re-entry.

Currently, spacecraft demise is assessed using low-fidelity engineering codes (object or spacecraft-oriented) that simulate satellite degradation along the entry trajectory. These tools rely on strong simplifying assumptions regarding heat flux prediction, ablation phenomena, and structural failure. Within the framework 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 by modeling the key aerothermodynamic, thermal, and structural phenomena that drive such events. This is achieved by coupling high-fidelity, physics-based numerical tools capable of simulating spacecraft aerothermodynamics in the rarefied and transitional regimes together with the thermo-structural behavior of the spacecraft structure.

The SMURFS toolset (Spacecraft Motion and behaviour Under Re-entry for Fragmentation Simulations) integrates trajectory, flow, and thermo-mechanical computations. Starting from a volumetric mesh of the spacecraft, the toolset progressively decomposes the geometry into fragments along the trajectory. A loosely coupled approach is adopted across the three modules: steady-state aerothermodynamic evaluations at predefined altitude stations are combined with transient 6-DoF (Degrees of Freedom) trajectory analyses and quasi-static thermo-mechanical responses. This framework enables fragmentation assessment through dedicated failure criteria — thermal criteria flag fragmentation when melting temperatures are reached, and mechanical properties are strongly degraded, while mechanical criteria, applied in a subsequent simulation step, detect fragmentation driven by the resulting stress states. When fragmentation occurs, each debris piece becomes independent and is propagated as a new simulation branch to compute its further decomposition.

This presentation provides a status update on the SMURFS toolset, summarizing recent developments, ongoing activities, and applications to representative test cases. We outline the main features of the toolset, discuss its underlying assumptions, and offer practical guidelines for setting up and running simulations. The maturity of the code has advanced significantly, culminating in the first official release to ESA, which includes a regression-test suite to ensure reliability across versions. Complementary developments include a graphical user interface (GUI) that streamlines case setup and execution, as well as a dedicated post-processing toolchain for visualizing and analyzing simulation results. We then describe mesh-erosion techniques that preserve the geometry's watertightness throughout the fragmentation process. Finally, we demonstrate the toolset's capabilities on a series of representative test cases, including multi-material CubeSats and the AVT reference model.

Speaker: Arnaud Francois (Cenaero)

28 Explosive Geometric Fragmentation and Collisions During Re-entry

With the ever increasing number of re-entries occurring year-on-year, accurate modelling of the entry process is a vital part of the drive for space sustainability. Both in terms of reducing the many unknowns of destructive re-entry and applications in satellite design according to the Design-For-Demise (D4D) paradigm. Of particular interest for D4D efforts is accurate resolution of explosive events as the dispersion of debris fragments caused by explosions can have a significant effect downstream casualty risk. Due to this interest, different methodologies for accounting for explosive events have been deployed by the re-entry simulation community. However, as a rule, they do not account for original object geometry.

The TransAtmospherIc flighT simulatioN tool (TITAN) is an open-source multifidelity tool that seeks to simulate re-entry across many computational domains, fidelity levels and modelling disciplines. In line with this logic and TITAN’s core functionality as a mesh-based tool it was considered desirable to explore fragment generation methods that incorporate information about original component geometry into the resultant fragments. Alongside this, recent work on the tool in areas of dynamic propagation and collision resolution have developed the capabilities necessary to model explosive events.

In this work a presentation of TITAN’s explosion model is provided. Fragments with complex geometry are generated after explosion through partitioning of the original component mesh according to a geometric statistics fragmentation approach. These individual fragments are then propagated as a debris cloud using adaptive time-stepping and binary search Time-of-Impact (ToI) prediction with multibody dynamics and interactions represented through an impulsive constraint-solving collision model.

After considering an exploratory simplified case, the model is applied to more realistic spacecraft re-entry contexts. An external case with an exposed tank such as would be seen on an upper stage rocket body is compared to an internal case where a tank explodes inside an enclosed bay. Downstream fragment dispersion is analysed as a result of such explosive events and ground impact footprint is also assessed.

Speaker: Tommy Williamson (University of Strathclyde Aerospace Centre for Excellence)

29 OSMOSE : Optimisation of Structural Morphologies for Operational Safety and End-of-life ; A topology optimization tool coupled to PAMPERO

Design-for-Demise methodologies are becoming increasingly important in spacecraft development as regulatory requirements and casualty risk constraints continue to evolve. In this context, structural components must simultaneously satisfy mechanical performance requirements during mission operations while maximizing their destructibility during atmospheric re-entry. This paper presents a newly developed topology optimization tool dedicated to demise-oriented spacecraft design.
Unlike conventional structural optimization approaches, where the objective function is primarily driven by mass reduction and stiffness performance, the proposed methodology incorporates additional geometric criteria specifically selected for their positive influence on demise behavior. These criteria include local curvature radius, thickness distribution, and geometric features promoting favorable shock-wave interactions and enhanced aerothermal loading during re-entry. By coupling structural optimization with demise-oriented metrics, the tool enables the generation of architectures that retain operational structural integrity while facilitating thermal degradation and fragmentation under re-entry conditions.
The developed framework combines thermo-mechanical constraints with aerothermal indicators derived from high-enthalpy flow analyses, allowing the optimization process to account for both in-flight mechanical requirements and destructive re-entry phenomena. Particular attention is given to the influence of localized geometric features on heat concentration mechanisms, shock interactions, and structural weakening processes.
The objectives of this work are twofold. The first is the development of a numerical design-assistance tool capable of supporting engineers during the early phases of spacecraft structural design. The second is the extraction of practical Design-for-Demise guidelines applicable to commonly used spacecraft structural elements. The study identifies geometric trends and design strategies that improve demise performance while maintaining acceptable structural efficiency.
Preliminary optimization results on brackets and bipod support structures demonstrate the potential of topology-driven demise design approaches to reduce surviving debris risk without major penalties on structural functionality. The proposed methodology represents a significant step toward the integration of demise-oriented criteria into future spacecraft structural optimization processes.

Speaker: Eddy Constant (R.Tech)

30 Analytical Analysis of the Re-entry Break-up Altitude Using Dimensional Analysis

Destructive atmospheric re-entry has become a major focus of current research, as it serves as the primary disposal strategy for satellites after their end of operation. The fragmentation of the satellite remains a key area of investigation, since its underlying mechanisms are not well understood. Observation data from multiple airborne observation campaigns conducted over the past decade show, as one key finding, the occurrence of a distinct explosive event triggering the further fragmentation process. We have named this event the main break-up and the corresponding altitude the main break-up altitude. Interestingly, the observation data show that the break-up occurs almost always at an altitude range from 75 km to 80 km for controlled as well as semi-controlled re-entries. This holds true for re-entries with various entry conditions and ballistic coefficients.
In this work, we apply a dimensional analysis to investigate the impact of the driving parameters on the break-up altitude. Based on this and under the assumption of comparable structural materials, a simplified model equation is derived
$h_\mathrm{Bu} = a \cdot \ln\left(\frac{v_0^4}{\sin(\gamma)}\right) + b$
The break-up altitude $h_\mathrm{Bu}$ is a function of the entry velocity $v_0$ and of the entry angle $\gamma$ only. The dimensional analysis reveals that characteristic parameters, such as the ballistic coefficient, play a minor role, which is consistent with the observation data. The model parameters a and b are identified by fitting the equation to data obtained from airborne observations. Additionally, numerical data were used, generated with the destructive re-entry tool SCARAB. The numerical data are used to account for uncontrolled re-entries, because there are no observation data available yet. The fitted function is in good agreement with both the experimental and numerical data, which indicates that the simplifying assumptions are reasonable and the simplified model equation captures the essential physics of the system. During the talk, we will present more insight into this approach and provide further discussion of the underlying simplifications.

Speaker: Clemens Mueller (HEFDiG, Institute of Space Systems, University of Stuttgart)

Discussion on the Update of the DIVE Guidelines: Feedback Collection

10:32 Coffee & Networking

Fragmentation modelling

31 Recent Modelling Improvements to the Spacecraft Demise Assessment Tool PAMPERO : PAMPERO HiFi

This paper presents recent developments implemented in the High Fidelity Branch of Pampero spacecraft demise assessment tool, aimed at improving the physical realism and predictive capability of re-entry survivability analyses.
The first major enhancement concerns the aerodynamic and aerothermal modeling framework. Previous implementations relied primarily on engineering correlations to estimate aerodynamic loads and convective heating during atmospheric re-entry. In the updated version of Pampero, these simplified approaches are supplemented by Computational Fluid Dynamics (CFD)-based methodologies, enabling higher-fidelity characterization of flow-field interactions, localized heating phenomena, and shape-dependent effects across a wider flight regime. The integration of CFD-derived data significantly improves the representation of complex geometries and transitional flow conditions.
A second area of improvement involves the mechanical response modeling of spacecraft components during fragmentation and structural failure. The upgraded framework introduces refined thermo-mechanical coupling. These developments provide a more realistic prediction of fragmentation sequences and surviving debris characteristics.
In addition, a thermite reaction model has been implemented to assess the influence of energetic material interactions on spacecraft demise behavior. The model enables simulation of exothermic reactions between selected metallic materials and oxidizers under re-entry conditions, allowing investigation of their potential to accelerate component destruction and reduce ground casualty risk.
The combined improvements substantially increase the fidelity of the Pampero tool while maintaining practical applicability for spacecraft design and certification activities. Preliminary validation and comparative analyses demonstrate improved agreement with high-fidelity reference cases and enhanced capability to evaluate advanced demise-oriented design concepts for future space systems.

Speaker: Valentin Ledermann (R.Tech Engineering)

32 Recent Modelling Improvements to the Object Oriented Tool DEBRISK : DEBRISK V4

The accurate prediction of spacecraft demise behavior during atmospheric re-entry is essential for assessing ground casualty risk and supporting Design-for-Demise strategies. This paper presents recent developments introduced in the CNES Debrisk software version 4, focusing on improved aerothermal fidelity through geometric shadowing effects and enhanced surface chemistry modeling.
A first major improvement concerns the implementation of object shading effects through the explicit definition of primitive positions and orientations within the spacecraft assembly. Previous modeling approaches assumed simple exposure conditions for individual components: only objects at the interior of other objects were full shielded from the flow. The upgraded Debrisk framework now incorporates the spatial configuration and attitude of geometrical primitives, enabling the computation of localized flow exposure and blockage effects during re-entry. This enhancement allows a more realistic estimation of heat flux distribution, temperature evolution, and component survivability, particularly for densely packed spacecraft configurations and complex geometries.
The second major novelty is the introduction of a catalycity model for surface heat transfer prediction. The implemented model accounts for catalytic recombination effects occurring at material surfaces under high-enthalpy atmospheric conditions, which can significantly influence convective heating levels during hypersonic flight. By incorporating material-dependent catalytic behavior into the aerothermal calculations, the updated methodology improves the prediction of wall heat fluxes and thermal response for a broad range of spacecraft materials.
Together, these developments substantially enhance the physical realism of the Debrisk software while preserving its applicability to engineering-level demise analyses. Preliminary assessments demonstrate the importance of both geometric shadowing and catalytic effects in determining component survivability and breakup behavior during atmospheric re-entry. The new capabilities provide improved support for spacecraft design optimization and future compliance with debris mitigation and casualty risk requirements.

Speaker: Martin Spel (RTECH)

Revisiting the VKI Free-Flying Ring Test Case

33 Revisiting the VKI Free-Flying Ring Test Case: A Possible Geometric Explanation for Drag Overprediction in CFD Rebuilding Efforts

The VKI free-flying ring test case has become an important benchmark for hypersonic shock-body interaction and six-degree-of-freedom free-flight prediction. Several independent CFD rebuilding efforts, including those from DLR, CIRA, Strathclyde, and RTECH, have reproduced the flow physics of the experiment but have consistently predicted a drag coefficient higher than that inferred from the experimental trajectory.
In the present work, simulations performed with the Mistral and Blizzard solvers revisit this discrepancy by examining the sensitivity of the result to the ring geometry and mass properties. A possible inconsistency was identified between the published model dimensions, the reported mass, and the expected mass obtained using the nominal aluminium density. In particular, the published ring thickness of 2.0 mm appears to lead to a mass larger than the reported experimental value.
As a working hypothesis, the ring thickness was reduced from 2.0 mm to 1.8 mm while keeping the rest of the experimental setup unchanged. This 10% reduction yields a substantially improved match with the measured free-flight trajectory and brings the predicted aerodynamic deceleration into close agreement with the experiment. The result suggests that the previously observed drag overprediction may not arise solely from modelling or numerical limitations, but could also be linked to uncertainty in the test-article geometry or mass properties.
This interpretation remains provisional: the proposed thickness correction has not yet been confirmed by VKI, and clarification of the exact manufactured geometry and material properties is still pending. Nevertheless, the study highlights the strong sensitivity of the flying-ring configuration to small geometric variations and underlines the need for precise, traceable mass-property data when using free-flight experiments as validation benchmarks.
The work further demonstrates the ability of the Mistral solver to reproduce complex hypersonic shock-interference dynamics when the reconstructed test-article properties are consistent with the observed motion.

Speaker: MARTIN SPEL (RTECH)

HIC SVNT DRACONES - re-enter with the dragon

Workshop Closure

Technical visit: Free visit of La Cité de l'Espace