Numerical and Experimental Investigation of a Two-Staged Supersonic Combustion Chamber

Abstract

Scramjet engines are expected to decrease the cost for access to space immensely. But due to manifold physical challenges, theses engines are far away from series production. Experimental studies are difficult to perform and flight tests are rather expensive. Numerical studies are, therefore, often used to perform parameter studies within the relevant Mach number regime. Nevertheless, systematic experimental data are absolutely essential when validating numerical studies. The present study focuses on a two-staged supersonic combustion chamber and hereby joins numerical studies with experimental validation and investigation. The two-staged combustor is expected to have essential advantages compared to the single-staged design. Previous studies investigating the singlestaged central injector concept at the Institute of Aerospace Thermodynamics (ITLR), Universitat Stuttgart, Germany, showed that the central injector bears high mixing capabilities, although the total amount of injected hydrogen is limited due to arising blockage effects. If intended to inject more fuel in order to increase the heat release and accompanied net trust a second injection stage generates the possibility to realize this by simultaneously avoiding the risk of high blockage or even flow separation in the combustor. This study, therefore, investigates a second injection concept, for a variation of equivalence ratios experimentally and numerically. The central injector for the first stage remains. The numerical investigations are validated with experimental data by means of wall pressure distributions and show reasonable agreement in their characteristics. Furthermore, we evaluate the performance of the second-stage concept by means performance parameters as stagnation pressure loss as well as mixing- and combustion efficiency.