Use of different oxidizers for afterburning of hydrogen heated by solar energy in rocket engine

Solar thermal propulsion (STP) with heating of hydrogen in high-temperature “concentrator-absorber” system with the possibility of its subsequent afterburning with cold oxidizer is considered. Such components as fluorine and hydrogen peroxide are suggested as oxidizers. These oxidizers create high-energetic propellant compounds with hydrogen and have high stoichiometric ratios, that allow to reduce required overall dimensions of solar energy mirror concentrator, as one of the most critical elements of STP. Results of parameters optimization of concentrator-absorber system by criterion of maximum of payload weight in the task of payload injection into geostationary orbit for the cases of isothermal (uniform) single-stage absorber and non-isothermal absorber with heating stages (non-uniform heating) use with reference to use of metal or film concentrator are presented. It is shown that requirements to accuracy of concentrator and its tracking system to the Sun are rather high. Decrease of dimensions of solar concentrator at afterburning of heated hydrogen simplifies its creation and improves operation in space conditions, in particular improves inertial characteristics of the STP as a part of space vehicle and simplify orientation to the Sun. Comparative characteristics of the STP use as space vehicle subsystem for the cases of metal and film concentrators are presented at isothermal and not isothermal heating of hydrogen and its afterburning by the considered oxidizers. The required area of concentrator in the case of fluorine use can be significantly reduced (up to 25 % and more) in comparison with oxygen-hydrogen STP. Use of hydrogen peroxide in some tasks allows 10…12 % reduction of the concentrator area as compared to oxygen-hydrogen STP. It is shown that payload injected into geostationary earth orbit, in the case of fluorine-hydrogen STP with film concentrator and non-uniform heating, can be increased up to 70 % and higher in comparison with the use of oxygen-hydrogen liquid propellant rocket engine.