Методика определения термопрочности топливного бака орбитальной ступени с маршевым жидкостным ракетным двигателем

Abstract

The present work is devoted to the development of engineering techniques for assessing the explosion of the space launch vehicle orbital stages, after turning off the main liquid propulsion engine, located in a circular orbit at altitudes of 200-1000 km. It is supposed that the unused residue of liquid propellant component in the propellant tank is vaporized by the impact on the structure of the pro¬pellant tank space factors, thus increasing the vapor pressure propellant to values exceeding the design strength of the pro¬pellant tank. Under the current outer space factors mean thermal effect on the orbital stage of the during its orbital motion (direct solar radiation, reflected from the Earth’s solar radiation, the Earth’s own radiation and aerodynamic heat flux). During the practical calculation it is solved a number of problems: the definition of the maximum (finding the orbital stage entirely in the lighted orbit) and the minimum (presence of maximum shaded portion of the orbit) limit of the thermal loading of the propellant orbital stage; the determination of the pressure of the vaporized propellant component vapor in the tank orbital stage (depending on the mass and the boundary conditions of the liquid propellant component placement residue); evaluation of strength tank design orbital stage by increasing its temperature and increased internal pressure caused by the evapora¬tion of the propellant component residue. To assess the explosion of the propellant tank orbital stage of the space launch vehicle with the main liquid propellant engine it is analysed the criteria in the form of ratios of ring voltage of a propellant tank structure to the value of the tensile strength, while the corresponding values of the lower and upper boun-daries of the thermal load. The calculations showed average values of absorbed heat flux at maximum thermal loading (the pro-pellant tank surface is exposed to the total direct solar radiation, reflected from the Earth’s solar radiation, intrinsic radiation of the Earth and aerodynamic heat flux throughout the orbit) and minimum thermal loading (the propellant tank surface orbital stage is in Earth’s shadow). The pre¬sence of residual unused liquid oxygen tank into the space launch vehicle “Zenit” in an amount up to 3 % of the initial filling of the tank does not contribute to exposure.