Determining Energetic Characteristics and Selecting Environmentally Friendly Components for Solid Rocket Propellants at the Early Stages of Design

This paper has investigated the possibility to theoretically calculate a value of the specific impulse for highly energetic compositions using only two parameters – the heat of the reaction and the number of moles of gaseous decomposition reaction products. Specific impulse is one of the most important energetic characteristics of rocket propellant. It demonstrates the level of achieving the value of engine thrust and propellant utilization efficiency. Determining the specific impulse experimentally is a complex task that requires meeting special conditions. For the stage of synthesis of new promising components, the comparative analysis of energetic characteristics, forecasting the value of specific impulse, especially relevant are calculation methods. Most of these methods were first developed to determine the energetic characteristics of explosives. Since explosives and rocket propellants in many cases have similar energy content and similar chemical composition, some estimation methods can be used to assess the specific impulse of solid rocket propellant. The specific impulse has been calculated for 45 compositions based on environmentally friendly oxidizers (ammonium dinitramide, hydrazinium nitroformate, hexanitrohexaazaisowurtzitane) and polymer binders polybutadiene with terminal hydroxyl groups, glycidylazide polymer, poly-3-nitratomethyl-3-methyloxetane). It was established that the estimation data obtained correlate well with literary data. Deviation of the derived values of the specific impulse from those reported in the literature is from 0.4 % to 1.8 %. The calculation results could be used for preliminary forecasting of energetic characteristics for highly energetic compositions, selecting the most promising components, as well as their ratios

Development of the Combined Method to De-orbit Space Objects Using an Electric Rocket Propulsion System

A method has been developed for the combined de-orbiting of large-size objects of space debris from low-Earth orbits using an electro-rocket propulsion system as an active de-orbiting means.A principal de-orbiting technique has been devised, which takes into consideration the patterns of using an electric rocket propulsion system in comparison with the sustainer rocket propulsion system.A procedure for determining the parameters of the de-orbiting scheme has been worked out, such as the minimum total speed and the time of the start of the de-orbiting process, which ensures its achievement. The proposed procedure takes into consideration the impact exerted on the process of the de-orbiting by the ballistic factor of the object, the height of the initial orbit, and the phase of solar activity at the time of the de-orbiting onset. The actual time constraints on battery discharge have been accounted for, as well as on battery charge duration, and active operation of the control system.The process of de-orbiting a large-size object of space debris has been simulated by using the combined method involving an electro-rocket propulsion system. The impact of the initial orbital altitude, ballistic coefficient, and the phase of solar activity on the energy costs of the de-orbiting process have been investigated. The dependences have been determined of the optimal values of a solar activity phase, in terms of energy costs, at the moment of the de-orbiting onset, and the total velocity, required to ensure the de-orbiting, on the altitude of the initial orbit and ballistic factor. These dependences are of practical interest in the tasks of designing the means of the combined de-orbiting involving an electric rocket propulsion system. The dependences of particular derivatives from the increment of a velocity pulse to the gain in the ballistic factor on the altitude of the initial orbit have been established. The use of these derivatives is also of practical interest to assess the effect of unfolding an aerodynamic sailing uni