Development of the concept of a reusable liquid rocket engine with three-component fuel

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Abstract

The article considers a promising direction for the development of liquid-propellant rocket engines (LPRE) – the use of three-component propulsion systems. The interest in this topic is based on a number of advantages that can be obtained by using this LPRE concept, namely: saving the mass of the launch vehicle (LV) by using a denser hydrocarbon fuel at the initial launch site; high specific impulse values at high-altitude launch sites due to the use of a more efficient pair of fuel components (FC): liquid oxygen + liquid hydrogen; reducing the cost of removing the payload, due to the use of a single propulsion system for both launch sites.

An analytical review of implemented three-component LPRE schemes developed in Russia and abroad has been conducted, and their main advantages and disadvantages have been highlighted.

Based on a detailed study of a number of circuit solutions for liquid-propellant rocket engines running on three-component fuel, the concept of a two-mode single-chamber three-component engine made according to a closed circuit with afterburning of generator gas is proposed. The oxidizer is liquid oxygen, the fuel is RG-1 kerosene and liquid hydrogen. In the first mode, the engine runs on three components, the share of liquid hydrogen in the fuel mixture is 4% of the total consumption of components. In the second mode, the engine runs on FC liquid oxygen + liquid hydrogen.

The results of a computational and analytical study of the optimal design parameters of the engine are presented. The aim of the study was to understand the qualitative picture of the influence of various fuel parameters on the thermodynamic properties of the combustion products of the fuel mixture and the engine efficiency. Based on the results of the study, the optimal percentage of fuel components was determined.

A mathematical model for calculating a three-component LPRE has been developed. The results of calculation of energy coupling are presented. A comparative analysis of the mass characteristics of the designed propulsion system is carried out.

About the authors

Vladislav A. Belyakov

Moscow Aviation Institute (National research university)

Email: titflavii@rambler.ru

Post-Graduate Student, Engineer of the Department 202 “Rocket Engines”

Russian Federation, 4, Volokolamskoe higway, A-80, GSP-3, Moscow, 125993

Dmitry O. Vasilevsky

Moscow Aviation Institute (National research university); Research and Testing Center of the Rocket and Space Industry

Author for correspondence.
Email: zudwa_dwesti_dwa@rambler.ru

Post-Graduate Student of the Department 202 “Rocket Engines”, Moscow Aviation Institute (National Research University); Engineer of the 1st Category, Federal State Enterprise “Research and Testing Center of the Rocket and Space Industry”

Russian Federation, 4, Volokolamskoe higway, A-80, GSP-3, Moscow, 125993; 9, Babushkina St., Peresvet, 141320

Alexey A. Ermashkevich

Moscow Aviation Institute (National research university)

Email: alex.ermashkevich@yandex.ru

Post-Graduate Student of the Department 202 “Rocket Engines”

Russian Federation, 4, Volokolamskoe higway, A-80, GSP-3, Moscow, 125993

Alexander I. Kolomentsev

Moscow Aviation Institute (National research university)

Email: a.i.kolomentsev@yandex.ru

Cand. Sc. Professor, Professor of Department 202 “Rocket Engines”

Russian Federation, 4, Volokolamskoe higway, A-80, GSP-3, Moscow, 125993

Ilnur R. Farizanov

AO “Ural Civil Aviation Plant”

Email: chelsea.physic@gmail.com

Design Engineer 1 Categories

Russian Federation, 1, p. 1, Marshal Zhukov Аv., Moscow, 123308

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Copyright (c) 2021 Belyakov V.A., Vasilevsky D.O., Ermashkevich A.A., Kolomentsev A.I., Farizanov I.R.

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