Научная электронная библиотека ULRICH'S Periodicals Directory

Сибирский государственный университет
науки и технологий имени академика М.Ф. Решетнева

Сибирский аэрокосмический журнал
ISSN 2712-8970 (Print) ISSN 2782-5760 (on-line)

Сведения об образовательной организации

UDK 629.78.03: 621.472 Doi: 10.31772/2587-6066-2019-20-2-251-265
SOLAR THERMAL PROPULSION SYSTEMS WITH VARIOUS HIGH-TEMPERATURE POWER SOURCES
Finogenov S. L., Kolomentsev A. I.
Moscow Aviation Institute (National Research University), 4, Volokolamskoe shosse, А-80, GSP-3, Moscow 125993 Russian Federation
The paper provides an overview of space thermal propulsion (STP) systems using concentrated solar energy as the main source of power. The paper considers solar thermal rocket engines of various configurations including those with afterburning of hydrogen heated in the “concentrator – absorber” system (CAS) with various oxidizers. Together with hydrogen the oxidizers form high-energy fuel compositions with a high value of ratio of components mass flow-rates which allows reducing the dimension of the CAS. The extreme dependences of the engine thrust on the specific impulse are shown for various values of the hydrogen heating temperature and the oxidizer-to-fuel ratio. The coefficients of the regression dependencies for the efficiency of a two-stage absorber and an absorber with the maximum non-isothermal heating having the highest possible energy efficiency are presented. The algorithms for calculating the main design parameters of the STP system as a part of a spacecraft (SC) are given, taking into account the ballistic parameters of the multi-turn transfer trajectory with multiple active segments applied to the STP systems having an energy-efficient non-isothermal CAS. The engine configurations with thermal heat accumulation and possible afterburning of heated hydrogen are also considered. Thermal accumulation allows accumulating energy in the solar-absorber during passive movement in the illuminated portions of the transfer orbits regardless of the lighting conditions of the apsidal orbit portions where the engine is turned on. Suitable heat-accumulating phase transition materials (HAM) such as the eutectic alloy of boron and silicon as well as refractory beryllium oxide are selected for different phases of the interorbital transfer to the geostationary Earth orbit (GEO). The main characteristics of different configurations of the STP systems in the problem of placing a spacecraft (SC) into high-energy GEO orbits are shown. A model of the SCSTP system operation is given taking into account ballistic parameters and the possibility of accumulating thermal energy. It is shown that the oxidizer-to-fuel ratio in STP systems with thermal energy storage (TES) increases with the decrease of the interorbital transfer time. The STP configurations with a two-stage TES showing a large energy-mass efficiency at moderate values of the solar concentrator accuracy parameter are considered.
Keywords: solar thermal propulsion, solar high-temperature heat source, concentrator-absorber system, thermal energy storage, hydrogen afterburning, ballistic efficiency.
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Finogenov Sergei Leonardovich – Senior researcher of Department 202 “Rocket Engines”; Moscow Aviation
Institute (National Research University). E-mail: sfmai2015@mail.ru.

Kolomentsev Alexander Ivanovich – Cand. Sc. Professor, Professor of Department 202 “Rocket Engines”,
Moscow Aviation Institute (National Research University). E-mail: a.i.kolomentsev@yandex.ru.


  SOLAR THERMAL PROPULSION SYSTEMS WITH VARIOUS HIGH-TEMPERATURE POWER SOURCES