Shock Tube Investigations of H2 Combustion in a High Temperature Supersonic Air Flow. (Scramjet)

Abstract

The experiments are performed in a constant area duct having a cross section of 60 mm by 60 mm. The shock tube generated flow has a stagnation enthalpy of about 5 MJ/kg (4200 K stagnation temperature) corresponding to the real flight of a hypersonic vehicle with Mach number 9 and at about 40 km flight altitude. The parameters of the flow at the entrance of the duct are approximately: flow velocity: 2500 m/s, static temperature: 1850 K and static pressure: 1.4 bar. By simultaneous monitoring of the evolution of both, the static pressures as well as the velocity profiles, a quantitative determination of the combustion efficiencies in the duct has been possible. In case of a combined injection mode, partly from holes in the sidewall and additionally from the top of small cylindrical tubes immersed into the flow under 30 degree angle, a rather effective mixing could be obtained leading to about 70% combustion efficiency 30 cm (5 times the height of the duct) downstream of the injection location.