Space Shuttle thermal protection system

The Space Shuttle thermal protection system (TPS) is the barrier that protected the Space Shuttle Orbiter during the searing heat of atmospheric reentry. A secondary goal was to protect from the heat and cold of space while in orbit. Materials The TPS covered essentially the entire orbiter surface, and consisted of seven different materials in varying locations based on amount of required heat protection:

Reinforced carbon–carbon (RCC), used in the nose cap, the chin area between the nose cap and nose landing gear doors, the arrowhead aft of the nose landing gear door, and the wing leading edges. Used where reentry temperature exceeded .
High-temperature reusable surface insulation (HRSI) tiles, used on the orbiter underside. Made of coated LI-900 silica ceramics. Used where reentry temperature was below 1,260 °C.
Fibrous refractory composite insulation (FRCI) tiles, used to provide improved strength, durability, resistance to coating cracking and weight redu

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

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

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