Ablation and radiation in hypervelocity earth-entry flows

Abstract

Atmospheric entry of spacecraft is an interesting and challenging problem in aerospace engineering as it involves a highly non-equilibrium aerothermodynamic shock layer environment. Re-entry of spacecraft into Earth's atmosphere occurs at very high entry speeds ranging from 8-12 km · s −1 , giving rise to a bow shock wave in front of the blunt body vehicle. The associated shock layer is characterised by a very rapid increase in pressure and temperature. The kinetic energy of the hypervelocity flow is converted into thermal and chemical energy, which is partially dissipated in the form of convective and radiative heat transfer. The gases traversing the shock layer undergo thermochemical changes such as dissociation, ionisation and recombination. To safeguard the re-entry vehicle from such a harsh thermal environment, a thermal protection system (TPS) is employed on the surface of the vehicle. The TPS material, when exposed to the re-entry heating, may undergo ablation due to the combined effect of heat flux and shear from the shock layer flow. The excited gas species in the shock and boundary layer react with the atomic/molecular species ablated from the solid TPS material. However, due to the uncertainties associated with the total heat flux estimation, large safety factors are currently being used in the TPS design, particularly for the afterbody region, compromising the payload mass and vehicle safety.The main objectives of this thesis were to experimentally simulate ablation product interactions with an expanding re-entry flowfield, to characterise the radiation of the entrained ablation products using ultra-violet emission spectroscopy targeting CN radicals, and to compare the ablating flowfields with non-ablating ones. The experiments were performed in the X2 expansion tube by using a stainless steel wedge model designed with a provision to mount a graphite ablation source on its compression face. The test flow condition, representative of a point in the Hayabusa capsule re-entry trajectory, was generated in X2 with a freestream velocity of 8.6 km · s −1 and a temperature of 2500 K, corresponding to an enthalpy of 38 MJ · kg −1 . As the test times available in X2 are limited, the ablating material cannot reach such wall temperatures by aero heating alone, and hence ablation in these experiments was created by electrically preheating the graphite strip to high wall temperatures representative of re-entry conditions. The wall temperatures realised in this work ranges from 1000-3000 K, which were measured by nonintrusive filtered image thermography using a dual-wavelength signal ratio technique. The hot graphite strip, upon exposure to the re-entry flow, ablates and mixes with the flow, and passes through the expansion fan and further into the afterbody region.The flowfield and interaction processes were optically diagnosed using a high frame rate video camera, two-dimensional filtered imaging and UV emission spectroscopy. CN emission measurements were made at various locations in the flowfield such as the fore...

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