Thermal Protection of the Huygens Probe During Titan Entry: Last Questions

CASSINI-HUYGENS mission is a cooperation between NASA and ESA, dedicated to the exploration of the Saturnian system. In the framework of this mission, the entry of the HUYGENS probe in the atmosphere of TITAN will be of major scientific interest. One of the essential points of the HUYGENS mission is therefore the good behavior of the thermal shield designed to maintain the aerodynamic shape and to protect the probe from excessive heating during the atmospheric entry on TITAN. The design and the qualification of this thermal shield were carried out between 1992 and 1995 (development phase). Currently, the final definition of mission parameters is being completed. As the performance of the thermal shield is one of all the parameters considered at system level, it is therefore necessary to reassess the thermal response of the TPS, taking into account some updated information that was not yet available during the development phase. After some recall of the results of 1992 to 1995, the paper will present a status of the current work on TPS

Breakup prediction under uncertainty: application to Upper Stage controlled reentries from GTO orbit

More and more human-made space objects re-enter the atmosphere, and yet the risk for human population remains often unknown because predicting their reentry trajectories is formidably complex. While falling back on Earth, the space object absorbs large amounts of thermal energy that affects its structural integrity.It undergoes strong aerodynamic forces that lead to one or several breakups. Breakup events have a critical influence on the rest of the trajectory are extremely challenging to predict and subject to uncertainties. In this work, we present an original model for robustly predicting the breakup of a reentering space object. This model is composed of a set of individual solvers that are coupled together such as each solver resolves a specific aspect of this multiphysics problem. This paper deals with two levels of uncertainties. The first level is the stochastic modelling of the breakup while the second level is the statistical characterization of the model input uncertainties. The framework provides robust estimates of the quantities of interest and quantitative sensitivity analysis. The objective is twofold: first to compute a robust estimate of the breakup distribution and secondly to identify the main uncertainties in the quantities of interest. Due to the significant computational cost, we use an efficient framework par-* Corresponding author. ticularly suited to multiple solver predictions for the uncertainty quantification analysis. Then, we illustrate the breakup model for the controlled reentry of an upper stage deorbited from a Geo Transfer Orbit (GTO), which is a classical Ariane mission

Infra-red and vibration tests of hybrid ablative/ceramic matrix technological breadboards for earth re-entry thermal protection systems

A new thermal protection system for atmospheric earth re-entry is proposed. This concept combines the advantages of both reusable and ablative materials to establish a new hybrid concept with advanced capabilities. The solution consists of the design and the integration of a dual shield resulting on the overlapping of an external thin ablative layer with a Ceramic Matrix Composite (CMC) thermo-structural core. This low density ablative material covers the relatively small heat peak encountered during re-entry the CMC is not able to bear. On the other hand the big advantage of the CMC based TPS is of great benefit which can deal with the high integral heat for the bigger time period of the re-entry. To verify the solution a whole testing plan is envisaged, which as part of it includes thermal shock test by infra-red heating (heating flux up to 1 MW/m2) and vibration test under launcher conditions (Volna and Ariane 5). Sub-scale tile samples (100×100 mm2) representative of the whole system (dual ablator/ceramic layers, insulation, stand-offs) are specifically designed, assembled and tested (including the integration of thermocouples). Both the thermal and the vibration test are analysed numerically by simulation tools using Finite Element Models. The experimental results are in good agreement with the expected calculated parameters and moreover the solution is qualified according to the specified requirements.European Comisión Fp7, 283797, HYDR

Astrium R&D Approach for TPS Development and Modeling

After a reminder of the main general considerations driving TPS development, this talk will present the current R&D project in which this subject is treated. An overview will be given about the scientific or industrial cooperations that are being implemented to address various topics such as material elaboration, material modeling, or related aerothermodynamic aspects

Group Discussions

This segment included: Imponderable and Webex call-ins Overview of Results Discussion of Next Round of Intercalibration Test Conditions by Alexandre Martin Future Validation Experiments (round table discussion) by Jean-Marc Bouilly Outcome of the Workshop and Future Directions by Ioana Cozmut

Uncertainty Quantification in Orbital Debris Reentry for Reliable Ground Footprint Estimation

The accurate prediction of the impact region of a reentering space object is a very complex problem. A typical reentry scenario involves three major steps. First, the object leaves its original orbit and starts falling back on Earth. This deorbitation step can take years in the case of a natural reentry or minutes in the case of a controlled deorbitation. The fragmentation step is likely to occur as the object enters more dense layers of the atmosphere, typically at an altitude of 100 kilometers. After fragmentation, parts of the object keep falling while being ablated by the atmosphere before hitting the ground or being entirely burnt by the atmosphere.The objective of this work is to accurately predict the impact region of the surviving fragments of an object. Due to the complexity of the physics involved during reentry, the scarcity of experiments and the unpredictability of events such as fragmentation, any model trying to predict a reentry features a limited reliability due to simplified assumptions and scarcity of data. In this work, we aim to alleviate this issue by coupling an original uncertainty quantification method with a space object reentry software developed by Airbus Safran Launchers in order to estimate the reliability of the numerical predictions. Although Uncertainty Quantification methods have been investigated for reentry problems but they could only include a limited amount of uncertain parameters due to the computational cost of classical uncertainty quantification methods in high dimension. In this work we propose a method that mitigates the curse of dimensionality and allows incorporating more uncertainties to the analysis. It exploits the fact that the solver making the predictions is a system of solvers. A set of interdependent solvers, where the output of an upstream solver is directly the input of a downstream solver), each of them solving a specific aspect of the physical problem. We first build a set of approximating functions of each solver based on global criteria and then they are assembled together in order to propagate the uncertainties through the system in an efficient way. Our mathematical framework allows the identification of specific solver, that affect the accuracy of the whole system. This feature allows to build adaptively an approximation of the global system and significantly reduce the computational cost. This method is then applied to the reentry problem under study

Intercalibration Results. Aerofast: Thermal/Ablation Analysis of the Front Heat Shield for a Martian Aerocapture Mission

An Aerocapture vehicle traveling from Earth to Mars, approaches that planet on an hyperbolic interplanetary trajectory. Upon arrival, the vehicle will perform a single atmospheric pass to significantly reduce its speed, and enters into an orbit around the planet. This maneuver uses aerodynamic drag instead of propulsion for orbit insertion, and potentially leads to large mass (fuel) savings as well as reduced flight times (higher arrival speed). However, Aerocapture results in significant aerodynamic heating, necessitating a Thermal Protection System (TPS), as well as the use of a guidance system to assure that the spacecraft leaves the planetary atmosphere on the correct trajectory. In the frame of the seventh European Community Framework Program (FP7), the AEROFAST (AERO- capture for Future space tranSporTation) research and development project aims at preparing a demonstration of a Martian Aerocapture mission and increasing the Technology Readiness Level (TRL). The aim of this poster is to present the preliminary thermo-mechanical analysis and design of the front-shield of the space probe. Despite the fact that several probe aerodynamic shapes and concepts are still being evaluated, this poster focuses only on the analysis of a 3.6 m diameter heat-shield, of an Apollo like shape with a low density phenolic impregnated cork as TPS. The 3D heat load history (convective and radiative), over the front-shield, is based on the maximum energy trajectory associated to a constant bank angle of 180◦ in a CO2 Martian atmosphere. During the first part of the project the implementation and validation of a 3D ablation and charring material model in the finite elements program SAMCEF Amaryllis was achieved. This was done by comparing the results of the 3D and 2D axisymmetric studies, performed on the leeward and the windward side of the probe. The numerical model consists of three sets of equations, namely the transient heat balance equation, the steady state mass balance equation and the charring equations. For the charring of the material we use a multi-species Arrhenius model with the species densities as degrees of freedom. The ablation is modeled by a surface imposed and temperature dependent ablation speed, followed by an in volume mesh deformation. During the second part, a preliminary design (allowable thickness versus curvilinear abscissa) of the TPS was obtained for the cork based material. The shape and mass evolution, during the Aerocapture phase, provides input for the mass budget and for the guidance and control analysis