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The 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. Additionally this year, the HyFAR-ARA association greatly contributed to the workshop as local organisation team.
As in all the previous editions, the upcoming Workshop has been designed with and for researchers, technicians, industry representatives, end-users and those with a general interest in the field, with the goal to exchange ideas, advance knowledge and discuss on the following topics:
The format of the event will be a sequential list of presentations of about 20 minutes each, with 5 minutes brief Q&A session per presentation.
REGISTRATION: CLOSED
Participation is free of charge, but the online registration is required.
CALL FOR ABSTRACT: CLOSED.
The timeline is the following:
16 September 2022: Registration deadline
16 September 2022 30 September 2022: Social Dinner registration deadline
16 September 2022: Abstract submission deadline
3 October 2022: Abstract selection announcement
10 October 2022: Complete program online
27-28 October 2022: Workshop
THE ATD3 WORKING GROUP
The ATD3 Working Group aims to achieve the following objectives:
On 27th October, a slot will be reserved for the Annual Meeting of the Working group members. All the WG members will be invited by e-mail.
Nowadays, besides the goal of enabling sustainable and more economical space access, the accumulation of inoperable objects is rising environmental and safety concerns. Various re-entry estimation software is being developed to promote the design of completely demisable spacecraft and hence mitigate the growth of space debris. Such toolkits perform dynamic analysis to compute the states of the re-entering objects, estimate the aerodynamic and aerothermal loads, study the fragmentation, and compute the demise and the ground footprint. These re-entry prediction tools are well suited for single bodies but usually have shortcomings in the analysis of the break-up and dispersion of fragments undergoing mutual interactions due to the complex physics involved.
This research aims to deepen the understanding of fragment interactions by supplying CFD and Design for Demise tools with data on such unaccounted proximal body interactions. In the present work, the aerodynamics of a free-flying ring model interfering with a steady crossflow cylinder (interstage rocket element and fuselage) in Mach 14 cold hypersonic flow has been investigated experimentally in the VKI Longshot facility. The ring model was initially suspended at $90^\mathrm{o}$ incidence upstream to a transversally mounted cylinder. Upon the arrival of the flow, the thin support wires of the ring rupture, allowing the model to fly freely and interfere with the two-dimensional shock generated by the stationary cylinder.
The measurement setup consisted of a set of freestream probes (Pitot pressure, static pressure, and heat-flux) and a schlieren system coupled with a high-speed camera. The transient freestream flow conditions have been computed from the probe data using an elaborate rebuilding technique. High-speed schlieren flow visualization has been performed to capture the motion of the model, and its trajectory and attitude were inferred using a synthetic image fitting algorithm. The analysis of the trajectory yielded the aerodynamic forces and moments, which were normalized with the free-stream dynamic pressure to compute the aerodynamic coefficients. A comparison with previous experimental data of Grossir et al. 2022 (https://doi.org/10.1007/s00348-020-02995-7) and modified Newtonian computations on an interference-free annular model, showed that the interaction with the detached shock wave vastly influences the flight trajectory. The interference induces a notable flownormal velocity even at early stages of the interaction promoting a shock wave surfing configuration.
The current experiment and the obtained results constitute the Test Case 2022 and the associated validation data for the Aerothermodynamics and Design for Demise (ATD3) Working Group.
The Navier Stokes Multi Block (NSMB) solver is developed in a consortium of different companies and universities in Switzerland, France and Germany. From the start the solver was designed to be used for hypersonic flow simulations, and it has been used in a large variety of ESA and EU funded space projects, among them the IXV (including the post-flight analysis) and the RETALT (Re-usable Launchers) projects.
To facilitate the mesh generation for complex geometries NSMB employs the chimera method was well as the patch grid method. For hypersonic flows a bow shock adaption procedure is available. A large variety of numerical schemes as well as turbulence models are available in the solver. The ALE (Arbitrary Lagrangian Eulerian) approach is available for moving grids as well as for moving bodies simulations.
Several CFD simulations were made for the ATD3 2022 test case. The grid for this test case was generated using the chimera approach with 2 groups. The first group includes the moving ring, while the second group includes the cylinder and serves as background grid for the moving ring. Particular attention was given to refine the background grid in the region in which it is expected that the ring will move. The grid was made such that 2 grid levels could be used in the simulations.
Calculations for the ATD3 test cases were made using the dual time stepping approach. The inflow conditions were varied in time according the test case description document. Calculations were made starting both at t=0 seconds and t=3.8 milliseconds. In the latter case the starting location of the ring as well as its initial state were given as input.
Simulations were made using different turbulence models, different outer timesteps, different tolerance criteria for the inner-loop convergence and different outer-time integration schemes. Most of the calculations were made on the coarse grid, but a comparison with the fine grid results is available for several cases. Initial results show a large influence of the selected turbulence model on the trajectory of the ring. The influence of the outer time step as well as of the outer-time integration scheme is small.
The DLR results of the ATD3 test case 2022 will be presented. The interaction of a free-flying ring with a two-dimensional curved shock wave has been simulated using the DLR TAU code. The flow solver was coupled with a 6-DoF motion module, dynamic grid adaptation has been used to capture the shock waves. A comparison between the numerical results and the experimental test case data will be presented and discussed.
Director of the Department "Sciences de l'Ingéniérie et du Numérique"
In the frame of the ATD3 working group, coordinated by ESA and CNES, a test case was proposed, representative of the interaction between two fragments generated by the break-up of a space vehicle during the re-entry.
The test was performed experimentally at the VKI Institute, and is described in detail in [1]. A ring is initially suspended over a cylinder (Figure 1), and when the flow enters the test chamber the wires are cut and the ring follows a trajectory over the cylinder, interacting with the shock wave generated by the cylinder itself.
Through the trajectory analysis, the aerodynamic coefficients for a number of different positions of the ring were calculated and are provided in [1]. One of the goals of the numerical test case is to simulate the ring in different positions and to compare the computed coefficients with the ones reconstructed from the experiment. An additional goal is to numerically compute the ring trajectory, by using the calculated aerodynamic coefficients.
In the present paper the focus will be on the first goal. All the conditions provided in [1] will be simulated, and the coefficients will be determined and compared with the experimental ones. If possible, the trajectory will be also calculated in a successive phase.
All the simulations are being performed with the NExT code, developed at CIRA. It is a multiblock structured code, able to simulate hypersonic flows including chemical and thermal nonequilibrium.
Since the code uses structured grids, a particular effort has been made in generating a topology that allows to deform the grid without changing the cells number and distribution for each block. Two strategies have been followed, by using on one side the grid generator ZENGRID, available at CIRA, and on the other side the commercial grid generator ICEMCFD: a sketch of the two approached strategies is reported in Figure 2. When ZENGRID is used, some of the features of the mesh may be parameterized to accommodate for several ring positions with respect to the cylinder and the computational grid can be then automatically generated providing in input the ring position and angle. In case of ICEMCFD, each grid is usually generated “manually” by the user, trying to optimize it for the specific ring position, even though a sort of automatization may be reached also in this case by using the ICEMCFD scripting language.
In Figure 3, a preliminary comparison is shown between the experimental values and the numerical results obtained by using the ZENGRID grid and the ICEMCFD grid. In figures 4 and 5 the Mach number contours for all the positions are shown for both the topology, even though with the ICEMCFD grid only a subset of the positions have been analyzed: in the final paper the full test matrix will be presented.
[1] D.G. Kovacs, G. Grossir, O. Chazot, “Hypersonic Aerodynamics of a Free-Flying Ring Interfering with a Two-Dimensional Curved Shock Wave—An Experimental Test Case” 2nd International Conference on Flight Vehicles, Aerothermodynamics and Re-entry Missions & Engineering (FAR) 19-23 June 2022.
Between 2000 and 2020, the number of man-made objects in orbit around the Earth has increased by approximately 82%, reaching a value close to 20000 objects, from which 53% are fragmentation debris. The current tendency is for the number of space objects to grow with the emergence of new satellite and CubeSat constellations.
To avoid the cluttering of space and decrease the risk of in-orbit collisions, they must be safely removed after reaching their end of life. A quite effective solution to this problem is to make the objects undergo destructive atmospheric re-entry, either controlled or uncontrolled, through which it breaks into several fragments, eventually demising due to the high aerothermal loads experienced during the re-entry process.
The accurate prediction of the destructive process and trajectory dynamics is of utmost importance to correctly assess the ground impact risks of surviving fragments. However, most state-of-the-art prediction tools use engineering and surrogate models that are unable to capture the occurring flow interactions formed by complex geometries and by the presence of multiple fragments in high-enthalpy regimes, which can impact the structural integrity and dynamics of the objects. Such flow features can be captured using high-fidelity tools, but their strict use is not computationally feasible.
To overcome this issue, the multi-fidelity based tool \textit{TITAN} (TransatmospherIc flighT simulAtioN) is being developed. The tool uses fully automated criteria to identify the level of fidelity required at each time step, enabling to switch between low-fidelity and high-fidelity models to compute the aerodynamic and aerothermodynamic quantities at the different flow regimes experienced by the bodies during the reentry process (e.g. rarefied, transitional, slip-flow and continuum regime), minimizing the uncertainty of the results. The dynamic motion of the objects is computed using the integrated 6 Degrees Of Freedom (DOF) trajectory propagator, enabling to analyse individual fragment trajectory.
This work investigates the importance of correctly resolving the generated flow features in predicting fragment dispersion due to the presence of multiple surrounding bodies. TITAN quasi-steady approach is used to numerically rebuild VKI's experiment of a free-flying ring crossing a shock-wave generated by a stationary cylinder in a Mach 14 Nitrogen flow, conducted at the VKI Longshot hypersonic tunnel. In addition, further validation of the high-fidelity model used within TITAN is conducted through the simulation of a standalone free-flying ring with several orientations and a comparison of the results with numerical and experimental data.
These 10 last years, the prediction of the space debris survivability during their re-entry and the associated prospective risk on ground are more and more in the scope of scientific research due to its complex multi-physics modeling and its crucial industrial applications, setting-up a permanent trade-off between fidelity of results and CPU costs. The use of the so-called ”high-fidelity” CFD method able to accurately model the forces and heat fluxes is a major challenge. One of those codes is MISTRAL-CFD, developed by R.Tech over the past 20 years. While originally developed for non-destructive reentry simulations, it has recently been applied to a large number of shapes used in object oriented demise tools (boxes, cylinders, rings) etc. While geometrically simple, those object represent very interesting flow features that are though from a numerical viewpoint (attached shocks, unsteady phenomena etc). Therefore, it is necessary to validate such codes before applying them to such shapes. A campaign on free flying (aero) and fixed shapes (heat fluxes and pressures) was carried out in the VKI financed by CNES, in order to assess the uncertainties in the simulations on such shapes (hollow hemispheres, rings). The use free flight method used in Longshot has been continued since, and one of the runs has been proposed as a test case for ATD3. The test case is focusing on the aerodynamics coefficients on a ring in interaction with a cylinder. MISTRAL has been used to rebuild the experiments, and the results will be compared in terms of aerodynamic coefficients over time, and flow visualization (Schlieren).
These 10 last years, the prediction of the space debris survivability during their re-entry and the associated prospective risk on ground are more and more in the scope of scientific research due to its complex multi-physics modeling and its crucial industrial applications, setting-up a permanent trade-off between fidelity of results and CPU costs. The use of the so-called ”high-fidelity” CFD method able to accurately model the forces and heat fluxes is a major challenge since it is necessary to carry out the meshing by hand, which can take days to weeks of engineering time for complex geometries.
Simplified aerothermodynamics tools based on empirical formulations are mainly used in order to reduce the computational cost of the simulation. Nevertheless, none of these approaches can estimate precisely the material degradation when complex physical phenomenon occurs, such as shock-shock interactions and wake interactions.
To resolve these issues, R.Tech(1), CNES(2), the French Defence Procurment Agency (DGA) and two French laboratories GEM(3) and M2P2(4) collaborate and develop a prototype using an automatic grid generation method based on octree Cartesian meshes [1] coupled to an Euler CFD solver. The major drawback of this type of approach comes from the quality of the near-wall modeling due to the nature of Cartesian meshes. Many methods have emerged since to overcome these problems as cut-cell, overset and Immersed Boundary Methods (IBM) (see [2], [3] and [4]), the last being use in this work. Although this is a great challenge and a medium-term perspective, the final objective would be to be able to replace the classical aerothermodynamics methods of PAMPERO by those developed in the prototype, in order to assess thermo-mechanical fragmentation.
This Euler solver is coupled to the finite element solver Code_Aster [5] in order to assess objects movements and deformations.
The present IBM is briefly described as well as the coupling scheme between Euler solver et Code_Aster. Von Karman Institute (VKI) Longshot free flight experiments on single objects and the ATD3 test case “Interference of free flying ring with a stationary cylinder” will be numerically rebuilt with both the new tools and the spacecraft-oriented tool PAMPERO. Then comparisons on the predicted free flight movement and the computation time will be performed.
References :
[1] S. Péron, T. Renaud, I. Mary, C. Benoit and M. Terracol, "An Immersed Boundary Method for preliminary design aerodynamic studies of complex configurations," in 23rd AIAA Computational Fluid Dynamics Conference, Denver, Colorado, USA, 2017.
[2] K. Luo, Z. Zhuang, J. Fan, N. Erl and L. Haugen, "A ghost-cellimmersed boundary method for simulations of heat transfer in compressibleflows under different boundary conditions," 2015.
[3] E. Constant, J. Favier, M. Meldi, P. Meliga and E. Serre, "An immersed boundary method in OpenFOAM : Verification and validation," Computers & Fluids, pp. Volume 157, pp 55-72, 2017.
[4] A. Pinelli, I. Naqavi, U. Piomelli and J. Favier, "Immersed-boundary methods for general finite-difference and finite-volume Navier-Stokes solvers," Journal of Computational Physics, 2010.
[5] EDF, "Code_Aster : Analysis of Structures and Thermomechanics for Studies & Research," [Online]. Available: https://www.code-aster.org/IMG/UPLOAD/DOC/Presentation/plaquette_aster_en.pdf.
Risk assessment of uncontrolled debris re-entering the atmosphere depends on various parameters among which drag and heat rates play a major role. However, those parameters cannot be computed with high fidelity methods such as CFD (Computational Fluid Dynamics) within a reasonable time frame for a full earth re-entry. Thus, correlations are usually used in spacecraft demise codes that use the object-oriented paradigm. However, correlations have often been derived in the 60s for non-destructive re-entry. A new methodology to compute drag forces and heat rates for destructive re-entry in the continuum regime, for a large range of geometry is presented. The models are based on CFD computations. The method is applied to complex shape for implementation in the object-oriented code DEBRISK v3. The model is extended to the transitional regime and its accuracy assessed thanks to Direct Simulation Monte Carlo (DSMC) computation. The DSMC code used is presented with its validation. Over 3400 simulations are carried out on 105 geometries in order to compute the random tumbling drag and heat rate in the transitional regime. The simulation setup to obtain reliable DSMC results in an automated way is outlined. Simulation results are compared with an approximation model. For the geometries investigated, parameters can be selected such that the average difference between the approximation model and the DSMC results are below 1% for both drag and heat rate
Risk assessment of uncontrolled debris re-entering the atmosphere depends on various parameters among which drag and heat rates play a major role. However, those parameters cannot be computed with high fidelity methods such as CFD (Computational Fluid Dynamics) within a reasonable time frame for a full earth re-entry. Thus, correlations are usually used in spacecraft demise codes that use the object-oriented paradigm. However, correlations have often been derived in the 60s for non-destructive re-entry. A new methodology to compute drag forces and heat rates for destructive re-entry in the continuum regime, for a large range of geometry is presented. The models are based on CFD computations. The method is applied to complex shape for implementation in the object-oriented code DEBRISK v3. The model is extended to the transitional regime and its accuracy assessed thanks to Direct Simulation Monte Carlo (DSMC) computation. The DSMC code used is presented with its validation. Over 3400 simulations are carried out on 105 geometries in order to compute the random tumbling drag and heat rate in the transitional regime. The simulation setup to obtain reliable DSMC results in an automated way is outlined. Simulation results are compared with an approximation model. For the geometries investigated, parameters can be selected such that the average difference between the approximation model and the DSMC results are below 1% for both drag and heat rate.
Simulating destructive re-entry is a demanding task that requires the modelling of non-equilibrium and high-temperature aerothermodynamics, structural and flight dynamics in presence of interacting shock waves, structural deformation, fragmentation and intense heat and mass transfer mechanisms. Three categories of methods can be identified to model re-entry: object-oriented, spacecraft-oriented and hybrid approaches that combine some characteristics of the former two. Existing low-fidelity methods based on semi-empirical or simplified analytical assumptions lead to substantial inaccuracies and uncertainties on fragmentation occurrences and trajectory paths due to their limits in properly accounting for the effects of highly nonlinear shock waves and structural responses during the destructive re-entry process. High-fidelity methods based on Computational Fluid Dynamics and Computational Structural Dynamics are becoming more and more capable of resolving the complex multi-physics of re-entry. Unfortunately, these methods bear a rather high computational cost that makes them unsuitable to perform repetitive simulations of different scenarios like those needed in the case of design for demise.
A recent trend, supported by the continuous improvement in computational capacity, is to bring together low- and high-fidelity tools to develop a collaborative and possibly a concurrent approach where a synergy between the methods is created and exploited. The exploitation of this synergy between high- and low-fidelity methods has recently led to the formulation and implementation of data-driven Reduced Order Modelling (ROM) based on Principal Components Analysis (PCA) and/or Manifold Learning or based on Physically-Informed Neural Networks (PINN) paired with autoencoders to keep the computational cost at a low level. Existing data, like high-fidelity simulations of a given problem, can be used to create ROMs that can incorporate the nonlinear physics and provide superior accuracy when compared to standard low-fidelity methods or physically-blind approaches such as Response Surface Methods (e.g. Kriging-based RSM or equivalent methods) while retaining high computational efficiency.
This work will serve to demonstrate the potential of this ROM approach for the prediction of aerothermodynamic loads for the Automated Transfer Vehicle (ATV) in the continuum regime. TITAN (the multi-fidelity framework developed during MIDGARD), SU2-NEMO (high-fidelity aerothermodynamics for continuum regimes) and RAZOR (adaptive ROM framework) will be used for this purpose, applying the PCA-based ROM Proper Orthogonal Decomposition (POD). Quantities such as Mach, Knudsen and object attitude and orientation will be applied as relevant parameters to define the re-entry corridor in this case. This work will therefore demonstrate the feasibility of such a ROM-based approach for the fast and accurate simulation of atmospheric re-entry.
The present study aims at proposing a methodology for coupling the non-equilibrium modelling solver(NEMO) of open source CFD software SU2 Multiphysics with the open source ablation solver Porous material Analysis Toolbox(PATO). SU2-NEMO solves Navier Stokes equations with thermochemical non-equilibrium effects by using finite volume method. Surface heat flux and pressure distribution of ablating wall is obtained from SU2-NEMO and it is used as input of ablation analysis. PATO, as a fully portable OpenFOAM library, discretizes conversation equations of total energy, gas momentum, gas mass, solid mass and gas species equations by finite volume method. As a boundary condition, PATO defines the convective heat flux with the formula $q_{conv}=u_e\rho_eC_h(h_r-h_w)$ where $u_e$ and $\rho_e$ are boundary layer edge velocity and density properties, $C_h$ is Stanton number, $h_r$ and $h_w$ represent recovery and wall enthalpy respectively. Therefore, to apply convective heat flux to the ablative surface, two different boundary conditions should be defined: multiplication of $u_e\rho_eC_h$ and recovery enthalpy($h_r$). Due to aerothermal design concerns, computational efficiency is significant for the coupling methodology. So that instead of estimating the boundary layer edge properties from CFD results of SU2-NEMO, which is computationally ineffective, shock relations are used for calculating $u_e$ and $\rho_e$ and recovery temperature formula is used for the calculation of $h_r$. Since $h_w$ is given as initial condition($c_pT_w$), $C_h$ can be calculated by using surface heat flux results. By using these boundary conditions PATO gives the outputs of surface temperature and recession which are used as inputs of SU2-NEMO for CFD analysis. This coupling methodology presents an effective way to investigate the effect of surface recession to the surface heat flux distribution of blunt nose geometries by using two open source solvers.
During atmospheric entry, capsules and space debris are exposed to a flow environment with complex fluid thermochemistry and gas-surface interactions (GSI) that may lead to mass loss and shape change. A promising approach for the numerical simulation of such challenging flows is the use of immersed boundary (IB) and adaptive mesh refinement (AMR) techniques, which offer reliable and efficient mesh generation and adaptation for moving shocks and recessing surfaces. For applications involving incremental geometry updates, such as objects reentering for demise, Cartesian grid based IB approaches substantially alleviate the strenuous task of generating adapted computational grids. Our project considers recent developments of such an IB-AMR solver based on a fully conservative cut-cell IB method able to incorporate GSI. A description of the methodology will be presented as well as updated verification and validation studies concerning the accurate modeling of thermochemical nonequilibrium, surface catalysis, and surface ablation. These results exhibit successful predictions in line with hypersonic simulations in literature concerning compression ramp and cylinder configurations, as well as with plasma wind tunnel experiments for a graphite ablator. Future direction for the development and unique potentials of the framework will also be discussed.
The present work aims at improving the numerical prediction of graphite material degradation during an entire reentry phase. For this purpose, 2D axisymmetric simulations are carried out on the nosetip of the IRV-2 vehicle, which is a well-referred test case that employed a thermal protection system composed of non-charring carbon. The coupled fluid / thermal approach adopted for such aerothermodynamics computations is presented, as well as two different ablation models, which both rely on the heterogeneous reactions of oxidation and sublimation that occur on the heat shield of the vehicle. The first ablation paradigm is based on the B’ tabulation, which has been historically used to assess the blowing rate of the recessing wall in the framework of a single gas (air) at chemical equilibrium. If such strategy has proven its reliability and efficiency over the past few decades, it suffers from many assumptions that are prone to be broken when applied to realistic descent trajectories. Such hypotheses include the consideration of a chemical equilibrium at the wall, a supposedly weak blowing rate, no injection of the ablated carbonaceous species into the flow, and the use of convective coefficients that directly depend on the boundary layer location. To overcome these aforementioned limitations, a more relevant ablation model, which takes care of the intrinsic multi-species nature of the flow, is proposed and implemented. The ablating mass flux and the species mass fractions at the wall, which are the product of finite-rate surface chemistry mechanisms, are accurately performed and updated at each convergence step of the flow. A particular emphasis is made on the generalization to chemical nonequilibrium, which leads to reducing the ablated surface thickness during a complete trajectory. In this perspective, the influence of the injected species that react with the surrounding flow is also investigated. Finally, the effects of different surface reaction schemes on the ablated surface shape and temperature are discussed.
Stijn Lemmens has been working for ESA’s Space Debris Office since 2011. His work has a strong focus on the analyses related to space debris mitigation and space traffic coordination, and their implementation in mission design, licensing, and operational processes in an international context. One of the major scientific problems in this field comes from the limited understanding of the physics associated with destructive re-entries. As such he is actively involved in the definition and running of projects to bring design for demise and break-up modelling from the scientists to the stakeholders.
During an atmospheric re-entry, a vehicle crosses at very high speeds the distinct atmospheric layers characterised by large density variations. The vehicle thus experiences several flow regimes, ranging from free molecular, rarefied (transition, slip) and continuous regimes. These regimes are commonly characterised by a Knudsen number (Kn) range [Kn is defined as the ratio of the mean free path and the vehicle’s characteristic dimension]. We focus on rarefied flows (0.001 < Kn < 0.01 for slip flow and 0.01 < Kn < 10 for transition flow). Such flows are mainly characterised by a non-zero slip velocity and a temperature jump in the vicinity of the wall as well as a diffuse aspect of the shocks. These flow features may strongly affect the aerodynamic properties of the vehicle compared to continuous flow. Besides, the Navier-Stokes equations - which assume small deviation from equilibrium - are no longer valid and a molecular approach is then required to describe such flows. Rarefied flows are therefore governed by the Boltzmann equation which takes into account both free flight and collisions of the particles. Simplified modelling strategies such as BGK-type models are also considered. These models are often solved using DSMC (Direct Simulation Monte Carlo) solvers such as SPARTA (Sandia National Laboratory), or deterministic discrete-ordinate solvers such as the CEA in-house K solver. It is crucial to ensure the validity of these numerical methods and modelling.
Since 2014, CEA-CESTA and CNRS/ICARE gather their respective expertise in experimental observations and numerical simulation/modelling of rarefied flows. On the one hand, this collaboration aims at taking profit of the experimental measurements for simulation validation and rarefied flow modelling purposes. On the other hand, the numerical simulations are used to complement the experimental observations. Therefore, the purpose of this presentation is to provide an overview of the latest studies performed in this framework, especially within the CHYP and APHYRA projects. As part of the CHYP project, measurements of Pitot pressure profiles, drag coefficient and glow discharge visualisations were performed at Mach 4 in the MARHy wind-tunnel around cone-cylinder geometries with 3 distinct aft-body (straight, rounded and flared) shapes. In the framework of the APHYRA project, measurements of aerodynamic forces and glow discharge visualisations were performed at Mach 4 and 20 in the same wind tunnel around a waverider model at several angle of attack. For both projects the main goal was to assess the influence of the rarefaction degree on the aerodynamic features. In the present work, numerical simulations for each geometry and flow condition with the CEA in-house K solver and SPARTA. In the presentation, comparisons between experimental and numerical results will be performed to assess the validity of the solvers for such various flow conditions and to illustrate the possible discrepancies between deterministic (K solver) and stochastic (SPARTA) approaches. Such an analysis is not commonly found in the literature. The simulation results will then be analysed to complement the experimental observations.
Composite materials like Carbon Fibers Reinforced Polymer (CFRP) are used in the manufacturing of components for satellites or launcher upper stages. Those materials behave similarly to thermal protection ablators hence show a strong resistance when exposed to high enthalpy flows. For instance, Composite Overwrapped Pressure Vessel (COPV) have been shown to survive the harsh atmospheric entry. This component is made of a metallic liner wrapped by CFRP. During the atmospheric entry, the part 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. High fidelity models for low density porous composite materials have been developed in recent years and their extension to CFRP is investigated in this paper. In particular, a unified numerical approach which solves the flow through and around the degrading porous material is considered. Volume averaging theory is used to described macroscopically the flow through the reactive porous medium and derive a single set of equations valid in the whole computational domain. This allows to capture with high accuracy the gas-surface interaction. This methodology has shown its advantages to predict the response of low density porous material but its extension to very dense composite material such as CFRP presents several numerical challenges that will be discussed. The high fidelity models are implemented in the high-order Discontinuous Galerkin code Argo which coupled with the Mutation++ library.
An experimental campaign is designed and performed inside the inductively coupled Plasmatron facility to validate the modeling and numerical prediction. Previous CFRP tests on coupons have shown that delamination of the fibers once the resin is completely pyrolyzed lead to overprediction of the demisability of the component. To avoid this phenomena, the sample is manufactured similarly to COPV. The scaled tank sample is then exposed to relevant atmospheric entry conditions in the Plasmatron facility. A first batch of CFRP tank shape sample have been tested and the material response reproduced numerically. The scaling down of real COPV leads to strong constraints for the manufacturing capabilities and affects the quality of the wrapping. A second batch of sample releasing those constraints is currently being manufactured and will be tested. The paper will present the experimental setup and measurements techniques for both campaign. The results of the experimental campaigns will be discussed and compared with preliminary numerical reproduction.
Among various techniques for Design for demise, maximizing the available heat to demisability is one of the complex problems which may require change in shape to increase the local heat flux, change in size to increase peak heat flux or adding additional energy by exothermic reactions. A frontal cavity has been studied in different applications and has been found to manipulate the heat transfers characteristics during hypersonic flights. A hemispherical shape can be considered one of the optimum shapes for demisability with high local heat flux on the edges and as demise progress from edges to center, it can further increase peak heat flux by decreasing ballistic coefficient. It also can provide the holding of exothermic heat addition if integrated with the system. However, frontal cavity in high-speed flows (like parachute or inflatable hypersonic / supersonic decelerator) can also have intermittent, non-stationary large amplitude, low frequency bow shock pulsations in front of it. During the absence of large amplitude fluctuations, small amplitude high frequency fluctuations are also observed in front of hemispherical cavity. Here large amplitude fluctuations can lead to reduced aerodynamic heating by emitting large vortices from the sides, while absence of these large amplitude fluctuations in frontal cavity can be favourable for demisability. In this experimental study, the effect of angle of attack have been studied on bow shock formation and its large and small amplitude instabilities in front of hemispherical cavity at hypersonic Mach number 7. The experiments for force measurements and Schlieren flow visualization with high-speed camera at 50000 frames per second, are carried out for different angle of attacks at Kashiwa Hypersonic & High-Enthalpy Wind Tunnel, The University of Tokyo. It is found that the bow shock formed in front of hemispherical cavity is subjected to nonlinear intermittent self-sustained fluctuations, which are quantified in time-domain using the image processing methods. The different bow-shock instability patterns are analyzed. Further it is found that these large amplitude fluctuations can be completely controlled by moderately changing the angle of attack of the cavity. The critical angle of attack after which these large amplitude fluctuations are completely controlled, are quantified in this study. This study only considers the aerodynamic effects in front of hemispherical cavity. In future it will be extended to aerodynamic heating analysis with and without direct hydrogen injection at the center of cavity.
The Plasma wind tunnel facilities at IRS have been used for ground testing since the 1980s. The basic techniques for material testing have remained largely unchanged. New developments and a renewed interest in the phenomena affecting and accompanying break-up of derelict satellites and space debris have necessitated the improvement of existing techniques, development of new techniques, or novel application of existing techniques.
We have followed a new idea for ground testing by introducing a mechanical load application method. Typical mechanical loads during re-entry have been characterized and may be applied during plasma wind tunnel testing. An extension of the LHTS technique of Kolesnikov has been proposed to correctly scale the plasma wind tunnel condition to the necessary geometries. This also requires novel heat flux measurements to characterize the conditions, which may be used to determine heat flux reduction for different materials. Spectroscopic measurements can be tailored to the intended application of the experiments, with the possibility to use a wide spectral range, or record specific spectral phenomena with kHz framerates. This data can be compared with airborne observation data collected during the last decades of observation campaigns. Non-intrusive photogrammetric measurements yield high-resolution 3-D surface information to track deformation as well as the formation of surface oxide layers.
The final presentation will present the application of new, improved, or novel application of techniques to comprehensively characterize material response during all phases of destructive re-entry.
The ISL hyperballistic tunnel is a hypersonic ballistic range. This unique facility in Europe [1] has been recommissioned in 2020 for hypersonic free flight and ablation studies. It is a combination of two two-stage light gas guns and of a 21 m long measurement tunnel.
The two light gas guns are used as model accelerators. Three launch tubes are available with calibres of 10 mm, 20 mm and 30 mm. By using hydrogen as light gas, the guns can accelerate models to velocities ranging from 2000 m/s to 9000 m/s corresponding to Mach numbers from 6 to 25. The measurement tunnel is isolated from the atmosphere, so that it can be fulfilled with different test gas selected in function of the study, typically nitrogen or air. Moreover, the pressure can be adjusted from a low pressure of 0.05 bar (h=20 km) to a high pressure of 2 bar. Consequently, the freestream Reynolds number ranges from 1E+4 to 1E+8 and thus covers the laminar, transitional and turbulent flow domains.
Ablation studies at ISL have started in the sixties [2]. Today, the experiments aim to extend the experimental database for ablation in high density flows with new geometries and new materials. The ablation measurements are achieved thanks to velocity measurements, high-speed imaging, X-ray radiographies and post-flight analysis. Recent ablation experiments of aluminium spheres were carried out with muzzle velocities up to 3500 m/s and have demonstrated the capabilities of the ISL hyperballistic tunnel for the ablation studies. The ablation on the sides of the stagnation points coupled with a light emission around the sphere and inside the wake flow were measured.
In conclusion, this hyperballistic tunnel is complementary to hypersonic wind tunnels, as some of their limitations can be overcome. For example, the projectile flies in a quiet test gas and the natural laminar-turbulent transition is reproduced. The modernization of the ISL hyperballistic tunnel is still an ongoing task and future ablation experiments will investigate the ablation of spheres made of different materials such as Inconel and titanium.
This study is conducted by the French-German Research Institute of Saint-Louis.
[1] E. Schülein, “Experimentelle Hyperschallversuchsanlagen und Messtechniken,” CCG-Seminar, Besonderheiten des Hyperschallflugs, 2019. (in German).
[2] J. Luneau, Contribution à l’étude des phénomènes aérothermiques liés au vol hypersonique décéléré. PhD thesis, Institut franco-allemand de recherches de Saint-Louis, 1969. (in French).
High-power lasers are increasingly considered for Active Debris Removal and illumination of low-orbit objects, and progress in high-power laser technology will be instrumental to achieve that aim. One of the main bottlenecks of high-power laser technology is the management of the ever-growing thermal load to which the optical elements are subjected. Scaling up their physical size indefinitely isn’t an option, as optical flatness is hard to attain for large surfaces. Moreover, the epitaxial growth of doped monocrystals often used in laser amplifiers is also limited in size, as dopant homogeneity becomes hard to control in large-diameter growth chambers. As a result, laser engineers have to come up with innovative laser architectures in order to maintain acceptable work temperatures as well as optimal optical quality required to focus lasers on low-orbital objects.
Based on the expertise of the CELIA lab in Bordeaux in high-power laser technology for scientific and industrial applications, we present a novel architecture for high-power pulsed lasers, developed in the HORIZON project. HORIZON’s main amplifier head is based on three active rotating discs fully immersed in ultra-pure cooling water. The water flow is controlled and maintained in a laminar regime, allowing the light to pass through and be amplified with negligible optical degradation. More than 1.3 kW have already been obtained on Horizon in continuous mode, and simulations predict pulses up to more than 800 mJ at 1 kHz, and a duration around 1 picosecond. The average output of HORIZON is therefore many times higher than the current systems used to test laser-assisted debris illumination.
Finally, we discuss the various innovations gathered to reach this goal, and conclude on the potential interest of Horizon-class lasers for the illumination and deorbitation of space debris.
The atmospheric reentry technologies have not evolved significantly since the early days of space exploration. Mostly they were designed for manned spaceflight applications allowing humans to return safely to Earth in a capsule or shuttle-like vehicle. Those applications have all the drawbacks that return vehicles whose shapes were constrained by launch phase restrictions (cross-section of the capsule or of the heat shield under fairing, aerodynamic shape of the shuttle). The next generation of reentry systems intends to disrupt this constraint with the deployment of a larger heatshield via an inflatable or deployable heatshield composed of flexible or textile thermal protection systems. If those concepts are not completely new and have already been intended in the past more than 2 decades ago, their implementation on operational applications seems to be finally underway. (H2020 EFESTO, NASA ADREPT, … ).
The trends for NEWSPACE and for GREENSPACE for a more industrial and sustainable use of space have reactivated the interest in reusable systems and recovery of elements from space in a more innovative manner (in-orbit manufacturing and production, short-time recovery of experiments out of ISS resupply cargo 6-month routine.).
Due to its experience in reentry topics coming from past activities in Design for Demise in a research entity of the ALTRAN Company, the e.NOVA start-up was created by his founder to promote this legacy and top listed this project BFS “Back from Space” as a main initial objective . The BFS project deals with an atmospheric reentry kit concept for Nanosat Payload based upon a aerothermal deployable heatshield with F-TPS (flexible Thermal protection system).
In that perspective, e.NOVA initiated a consortium with industrial textile developer, shape memory alloy provider, thermo-mechanical and computational fluid dynamics analysts.
This paper intends to present the state of the art on this thematic and the initial progress status of this internal project and the legacy reused in this initiative. It will introduce the upcoming activity related to F-TPS development, characterisation and validation. This activity has been identified as an “Innovative Technology for Nanosat” in the scope of French “Plan de Relance” aiming to address the Newspace market in a low-cost and mass production approach for in-orbit servicing and manufacturing in orbit. The BFS project aims to close the loop for Nanosat operators to be able to carry autonomously payloads back from orbit.