An experimental investigation of a Mach 2 combustor has been conducted in order to characterize flow properties in a supersonic reacting flowfield. Hydrogen was injected transversely as staged, underexpanded jets behind a rearward-facing step into a ducted Mach 2 air freestream. The effects of the chemical reaction on the supersonic flowfield was investigated using shadowgraphs, broadband flame emission photography, and planar laser-induced fluorescence of OH. The shadowgraphs indicated that the wave pattern in the combustor along with flowfield unsteadiness was strongly affected by the heat release. The broadband flame emission photographs revealed large regions of no combustion in the vicinity of the fuel injectors where fuel/air mixing was insufficient to support combustion. These regions decreased in size as the freestream stagnation temperature was decreased for fixed hydrogen mass flow rate, consistent with an increase in the effective g-ratio with combustion. The size of the zones containing OH in the planar fluorescence images also increased as the main flow stagnation temperature was decreased. Reaction zones were found in the planar fluorescence images away from regions containing inject ant in a nonreacting study of the same geometry, indicating that the pressure rise associated with the reaction forced a large redistribution of the fuel.
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