Numerical Simulations of the Apollo 4 Re-entry Trajectory

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.

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

Ground to Flight Investigations of Hayabusa with Ablation Effects

Nikhil Banerji, Elise Joy Fahy, Pénélope Leyland, Valentin Marguet, Jérémy Raphaël Mora-Monteros, Daniel Potter

Thermal protection systems (TPS) are of extreme importance for the survival of space vehicles especially during superorbital re-entry to Earth. The design of thermal protection systems requires in-depth knowledge of the thermal loading experienced during a re-entry. The thermal loading data is mostly determined using ground testing and can be backed up by modelling activities including computational fluid dynamics (CFD). The verification of this data with flight data is invaluable and a recent, rare example of an opportunity for comparison was the Hayabusa asteroid sample return mission, which landed in Australia in 2010. During this re-entry a team of international scientists collected spectral data which can now be used for comparison and verification of ground test and modelling/CFD data. Ground testing of subscale models at flight equivalent hypervelocity flow conditions (8 − 12 km/s) can be performed in hypersonic impulse facilities such as the X2 expansion tunnel at The University of Queensland. A recently developed method at The University of Queensland enables heated reinforced carbon-carbon (RCC) models to be tested at temperatures repre- sentative of those experienced in flight ( 2000 − 3000 K), in addition to testing with cold-wall metallic models. Hot wall testing allows more realistic simulation of re-entry flow characteristics including important thermal surface effects (surface chemistry, catalycity) which has previously not been possible. The eilmer3 compressible flow CFD code is used extensively for simulating atmospheric re-entry vehicles at flight and laboratory conditions. Simulations of the Hayabusa aeroshell incorporate heated walls, as well as surface catalycity, to accurately model the conditions experienced by the TPS. These simulations can be coupled with SACRAM, a 1D ablation modelling code, to include the effects of ablation and recession at critical points on the model surface. Current work is investigating the effects and validity of heat flux scaling correlations applied to a range of scaled models with the Hayabusa geometry and flight equivalent flow conditions. This will be achieved through results of CFD simulations, incorporating radiation and ablation modelling, and expansion tunnel testing with hot and cold wall models. Increased understand- ing of scaling methods will allow higher fidelity heat loading data to be acquired allowing more efficient and effective design of TPS. This paper will discuss results from CFD simulations coupled with ablation modelling and modelled surface chemistry. An outline of planned experiments in the X2 expansion tunnel, including background on the RCC heating method and condition development, will also be presented.

2013

Fluid-structure thermal coupling and ablation effects in atmospheric entry

Ojas Joshi

EPFL2013

Ground to Flight Investigations of Hayabusa with Ablation Effects

Nikhil Banerji, Elise Joy Fahy, Pénélope Leyland, Valentin Marguet, Jérémy Raphaël Mora-Monteros, Daniel Potter

2013