Optimization of scramjet inlet based on temperature and Mach number of supersonic combustion

Abstract

A two-dimensional mixed compression scramjet inlet design is presented in this work. The design is based on the static temperature and Mach number at the combustion chamber inlet according to the conditions required to burn hydrogen spontaneously at supersonic speed. This method considers air as a calorically perfect gas, with no viscous effects, and shock on-lip and shock on-corner. To burn hydrogen spontaneously at supersonic speed the mixture temperature of the income airflow and hydrogen is obtained considering the zeroth and first Laws of Thermodynamics. The criterion of equal shock strength, based on the normal component of the airflow velocity approaching the incident oblique shock waves, is applied to obtain the compression ramps angles and the airflow corresponding thermodynamic properties. Therefore, the total pressure ratio across all incident oblique shock waves and total temperature at the compression section are constants. Although the present method can be applied to any scramjet with a mixed compression system for any number of ramps, cases with up to 5 ramps, flying at a hypersonic speed corresponding to Mach number 7 through the Earth's atmosphere at 30 km of geometric altitude are considered. Finally, the same optimization criteria are extended to analyze the inlet of a scramjet vehicle with 5 compression ramps, flying at speeds from Mach numbers 5 to 10, at an altitude of 30 km.