Tte stage deorbiting with a deceleration pulse detonation engine

Рубрика: 
Space Energy, Power and Propulsion
1Zolotko, OE, 1Zolotko, OV, 1Sosnovska, OV, 1Aksyonov, OS, 1Savchenko, IS
1Oles Honchar Dnipro National University, Dnipro, Ukraine
Space Sci. & Technol. 2021, 27 ;(4):32-41
https://doi.org/10.15407/knit2021.04.032
Язык публикации: Ukrainian
Аннотация: 

The article discusses the issues related to reducing the amount of space debris from rocket stages. The main ways to remove the separable part of a rocket from a space orbit are: the usе of a deceleration detonation propulsion system; gasification of fuel residues and the use of a gas-reactive deceleration pulse system; continuation of the work of the main propulsion system after the separation of stages; the use of a harpoon to capture the rocket stage and the use of sail for its further braking; the use of anti-missile or combat lasers to destroy a stage on the orbit followed by the stage fragments’ burning in the Earth’s atmosphere.
       To select the optimal method for removing from the orbit the separated part of a space rocket, the arithmetic progression method was applied. It has certain advantages over the classical hierarchy analysis method and has no inherent disadvantages of this method. A ranked row of solutions was obtained according to the five most significant performance criteria, and its stability was proved.
      A new deceleration detonation propulsion design scheme is proposed. Detonation burning of residual fuel components provides the maximum possible value of the deceleration thrust impulse. Using the example of the second stage of the “Zenit” launch vehicle, we analyzed the nature of the dependence of the entry angle into the atmosphere on the important characteristic parameters: the deceleration speed impulse, the entry speed into the Earth’s atmosphere of the separated launch vehicle stage, the required value of the specific thrust impulse of the deceleration propulsion system. A new analytical formula has been obtained, which connects the thrust and specific thrust impulse values of the detonation engine with the determined detonation process parameters. The results of the computational experiment were compared with the results of calculating the specific thrust impulse using the new formula for oxygen-based fuel compositions, known experimental data, and numerical simulation data of other authors. The data obtained in this study make it possible to evaluate the design parameters of the deceleration detonation engine at the stage of analyzing technical proposals.
Ключевые слова: separable part of a rocket; deceleration detonation propulsion system; deceleration detonation propulsion design scheme; the arithmetic progression method; specific thrust impulse
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