Rentrée atmosphérique | EPFL Graph Search

vignette|Entrée atmosphérique de Mars Exploration Rover (MER), vue d'artiste. vignette|Représentation par un ordinateur de la rentrée atmosphérique de la Navette spatiale américaine. La rentrée atmosphérique est la phase durant laquelle un objet artificiel pénètre dans l'atmosphère d'un objet céleste et atteint des couches suffisamment denses pour provoquer des effets mécaniques et thermiques. L'objet en question peut être satellite, capsule spatiale ou fragment de fusée par exemple. Les objets conçus par l'Homme Le domaine est apparu dans les années 1950 aux États-Unis et en URSS, d'abord pour des objectifs stratégiques, concernant les missiles à longue portée disposant de véhicules de rentrée, puis pour un objectif politique : l'homme dans l'espace. D'autres nations ont suivi et la connaissance du domaine est assez mondialement partagée sur le plan scientifique, même si peu de pays disposent aujourd'hui d'un lanceur capable de réaliser un vol humain : la Russie, la Chine

Numerical Simulations of the Apollo 4 Re-entry Trajectory

Ermioni Papadopoulou

Sample return capsules, as the Apollo Command Module have been widely used to ad- vance the knowledge and planning of manned lunar and planetary return missions. Such reentry vehicles undergo extreme thermal conditions, caused by shock-heated air during their super-orbital atmospheric re-entry. Such extreme conditions can result in failure of the aeroshell structure and loss of important payload. This technological challenge is ad- dressed by the use of ablative thermal protection systems (TPS), which dissipate the heat away from the vehicles front wall via ablative products release into the boundary layer. Additionally, such velocity and temperature magnitudes during reentry conditions intro- duce significant radiative heat loads, filling the shock layer with radiators that react with the ablative species injected by the capsule wall. Therefore, accurate numerical modeling techniques are required, so that the thermophys- ical, thermochemical environment of a reentry capsule can be successfully reproduced and predicted. The present work aims to numerically rebuild certain significant trajectory points, containing the peak heating points of the Apollo 4 terrestrial re-entry. This re- quires the coupling of the resolved flow-field with radiative and ablative effects in order to accurately predict the convective and radiative heat flux for each trajectory point. The results will be compared to previous calculations and existing flight data. The numerical simulations are performed in 2D thermal non-equilibrium with a compress- ible explicit Navier-Stokes solver, coupled to a radiation database and a thermal material response code to implement the ablative effects. The calculations are performed also in 3D, using a commercial implicit Navier-Stokes solver. The results will be used to reproduce the capsules trajectory and verify the accuracy of the associated Modeling Tools.

2014

Ablation-Radiation Coupling Modelling for Hayabusa Re-entry Vehicle

Valentin Marguet

Reentry vehicles undergo extreme thermal conditions as they reach hypersonic velocities in particular conditions. Thus thermal protection system (TPS) are required to prevent the probe to be damaged. When it comes to Earth’s reentry, capsules like Hayabusa are equipped with an carbon phenolic TPS which ablates and releases ablation products into the boundary layer during reentry. Besides, as radiation can be a significant component of the overall heat load, chemical reactions may occur between the radiators and the ablative injected species so that the whole aerothermodynamic state is modified. Therefore, an accurate numerical modeling of the thermal, physical and chemical processes induced by hypersonic reentry , is needed. It implies to couple the fluid solver and the material response code thanks to a partitioned coupling algorithm. As for Hayabusa’s reentry simulations performed for the ARC project, ablation-flowfield and radiation-flowfield resulted to be accurate as they were compared to flight data or empirical correlations. Next step of ablation-radiation coupling consists in experimental tests in plasma wind tunnels and full coupling algorithm.

2013

Fluid-structure thermal coupling and ablation effects in atmospheric entry

Ojas Joshi

This thesis deals with the modelling of fundamental aspects of the mission engineering, the physics, and the constraints of atmospheric entry of space vehicles. The aim of this work is to formulate a coupled approach for the modelling and the simulation of thermal fluid-structure interaction in atmospheric entries, taking into account the thermochemical and physical response of the material composing the solid structure. Numerous missions that aim to explore planets, comets or regions outside our galaxy are planned for the next few years. During the phase of entry into the atmosphere of a planet or a moon, space vehicles encounter extreme heating conditions that can damage the structure and the on-board equipment. A Thermal Protection System (TPS) is hence installed to insulate the vehicle’s internal parts from high temperatures and heat fluxes. Ablative materials, which are composed of high-temperature resistant fibre porous structure filled with an organic matrix, are often used to manufacture heat shields of high speed entry vehicles. The hot gas around the spacecraft transfers energy fluxes to the vehicle through the fluid-structure interface. The solid then reacts, with part of the heat being conducted in the deeper layers of the material, and part being absorbed by physical and chemical transformations of the material itself. The organic phase of the ablative material evaporates – this process is called pyrolysis – and these gases are blown into the outer flow, creating thus a protective layer. Moreover, the surface of the TPS may react with the atmospheric gases, consuming in this way a fraction of the material. A model for these two phenomena and the thermal fluid-structure interaction is proposed in this thesis. Numerical simulations are one of the most widely used tools to design the TPS, and a detailed mathematical model allows to reduce margins in the sizing. Atmosphere gases and the spacecraft structure form two distinct systems that interact through the external surface of the TPS, and through which energy and mass exchanges occur. Many computational tools exist to simulate the physics of each of the two domains. A higher level of accuracy is however achieved if the problem is approached in a coupled fashion, where each system takes into account the thermophysical response of the other. An innovative coupling algorithm, which includes the direct evaluation of the surface chemistry, is presented in this thesis. Using this coupling procedure, numerical simulations of the TPS of a European Space Agency project are performed and key parameters for the sizing of the heat shield are evaluated and discussed.

EPFL2013

Fluid-structure thermal coupling and ablation effects in atmospheric entry

Ojas Joshi

EPFL2013