Solar thermal rocket engine with beryllium-oxide phase-transition latent heat energy storage and hydrogen afterburning

Аuthors

Finogenov S. L.

Moscow Aviation Institute (National Research University), 4, Volokolamskoe shosse, Moscow, А-80, GSP-3, 125993, Russia

e-mail: sfmai2015@mail.ru

Abstract

The article considers solar thermal propulsion (STP) with thermal energy storage (TES) containing high-temperature phase-transition material – beryllium oxide possessing high latent heat of phase transition “fusion-crystallization”. High melting temperature allows obtain the engine specific impulse at a level of 9000 m/s.

Joint optimization of the basic relevant parameters, such as masses ratio of solar mirror concentrator and TES in combination with mirror accuracy parameter was performed. It was demonstrated that that the ratio of TES energy capacity to solar radiation receiver thermal power, or ratio TES energy capacity to solar concentrator area in conjunction with optimal selection of accuracy parameter of the mirror can be accepted as an optimizing parameter. Maximum mass of a spacecraft being placed into geostationary orbit with time limitation of inter-orbital transfer from 30 to 90 days was selected as optimization criterion. Optimization was performed out by Gauss-Seidel method.

The optimization results revealed that optimal ratio of TES energy capacity and light detector power was 22-24 MJ/kW, which corresponds to the optimal ratio TES energy capacity to the concentrator area of 6-7 MJ/m² at rational mirror accuracy parameter of 0.25 degrees. The STP characteristics with TES are presented and analyzed. The article shows that for relatively small flight time of 3040 days optimal values of excess oxidant ratio corresponding to payload mass maximum. The higher value of excess oxidant ratio corresponds herewith to the lower value of the flight time.

Dependences of the TES energy capacity and the concentrator diameter from excess oxidant ratio for a wide interval of flight duration are presented. Expedient areas of heated hydrogen afterburning application for various inter orbital flight duration were determined. The article shows that afterburning is expedient for the time of putting to geostationary orbit of 30 to 45 days. The corresponding excess oxidant ratio changes herewith from 0.3 to 0.1. For the flight above 50 days, the monopropellant hydrogen STP is expedient. Compared to alternative inter-orbit transportation means, employing the combination of small and large thrust engines combination, the gain is about 450 kg under the one and the same inter-orbital transportation time of 60 days.

Keywords:

solar thermal rocket engine, latent heat thermal energy storage, beryllium oxide, hydrogen afterburning, geostationary orbit

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