Effect of bipropellant combustion products on the rocket nozzle design

Abstract

The focus of this research work is to investigate numerically the effect of adding the gas on the design and performance of axisymmetric MLN nozzles. A FORTRAN code was developed to design this nozzle using the characteristics method (MOC) at high temperature. The thermochemical and combustion studies of the most used liquid propellants on the satellites and launch vehicles allow to known all gases. Four engines are investigated: Ariane 5 (Vulcain 2), Ariane-5 upper stage engine (Aestus), Zenit first stage (RD-170) and Falcon 9 upper stage (Raptor). Thermodynamic analysis of parameters design MLN (such as length, Mach number, mass, thrust coefficient) was conducted. The comparison shows that the presence of 50% of H2O gas in combustion species increases the nozzle design parameters (diatomic gas including air) in the order of 25%. On the other hand, the existence of CO2 gas considerably increases approximately 35% the length and the exhaust radius. These rise depend on gases percentage in the combustion. The truncation method is applied in the MLN nozzles to optimize the thrust/weight ratio.