A Hypersonic Flow Effect on the Melting Rate of a Heat-Protective Surface under Degradation Conditions
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Vol. 61, No. 1, May, 2020
HEAT ENGINEERING A HYPERSONIC FLOW EFFECT ON THE MELTING RATE OF A HEAT-PROTECTIVE SURFACE UNDER DEGRADATION CONDITIONS N. I. Sidnyaev1,2 and E. V. Belkina1 Translated from Novye Ogneupory, No. 1, pp. 20 – 27, January, 2020.
Original article submitted October 11, 2019. The paper presents the results obtained for the distribution of heat flux incident upon a refractory plate of a hypersonic flight vehicle moving at space velocities at various distances from the Earth’s surface. The paper also provides the results of studying phase transitions in the near-wall boundary layer occurring during the hypersonic flow over the ablating surface. The effect of the catalytic wall on heat flux is analyzed. The emphasis is on the analysis of mass loss from the surface of high-velocity flight vehicles based on detailed consideration of the mechanism of heterogeneous catalytic reactions taking place under the surface mass transfer conditions. A temperature distribution is provided across the thickness of the boundary layer at the stagnation point of a blunt-nosed body with a refractory coating for a specific section of the flight path. The mass loss from the surface of crystalline refractory bodies is determined. Keywords: hypersonic flow, melting, mass loss, heat-protective surface, refractory tile.
elastomer film is applied to its surface prior to installation. To protect flexible joints, the outer layer of the coating is made of special ceramic fibers. In this case, protected units can simply be wrapped with such coating. The surface segments exposed to 371 – 649°C are protected with a coating consisting of 99.7 % of pure amorphous fire-resistant quartz fiber with colloidal silicon dioxide added as a binder. The coating is fabricated in the form of refractory tiles measuring 203 ´ 203 mm, which are 5 to 25.4 mm thick. The external surface of the tiles is coated with borosilicate glass containing a special pigment (SiO2-based white coating and Al2O3-based reflective coating) to ensure low absorptivity of solar radiation and high emissivity. Thermal protection of the body segments exposed to 649 – 1,260°C is realized by means of a reusable thermal insulation; the difference consists in the tile sizes (152 ´ 152 mm with a thickness varying from 19 to 64 mm). The spacecraft nosecap and the leading-edge assembly exposed to temperatures in excess of 1,600°C are protected with a refractory carbon-based material reinforced with carbon fibers. As the vehicle returns back to Earth, this material is degraded (melted) and must be replaced with a new one before each subsequent flight.
INTRODUCTION Studying heat transfer in a semi-infinite body, the surface of which is degraded at high temperatures as each kilogram of lost mass absorbs a certain amount of heat, is quite important. Despite an idealized statement of the problem, the existing techniques [1 – 6] contain all the main features of a non-steady state degradation of the actual heat-protective coatings, and are especially convenient when performing bench
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