High pressure LOx/H ₂ combustion and flame dynamics preliminary results

A windowed liquid rocket engine thrust chamber (designated BKC) has been operated over a broad range of conditions at reduced pressures (P of approximately 0.8, 1 and 1.2 with respect to the thermodynamic critical pressure of oxygen. Liquid oxygen (LOx) and gaseous hydrogen (G3/4 ) have been injected through a single coaxial injector element at temperatures of ca. 120K and ca. 130K-270K respectively. The coaxial injector geometry has been modified to provide a broad range of operating conditions in terms of propellant injection velocity ratio (V _R) and propellant momentum flux ratio (J). Values of V _Rrange from approximately 3 to 70 and J from around 0.08 to 20. An advanced measurement system has been developed to provide high accuracy propellant pressure and temperature measurements for determination of precise injection conditions. High speed optical diagnostics have been implemented including high speed emission imaging of H ₂0 and OH chemiluminescence in the near injector region at frequencies up to 9kHz. Shadowgraph imaging has been performed at up to 4kHz to visualize the flowfield up to 100 mm downstream from the injection plane. Preliminary results are presented here highlighting flame stabilization and flowfield characteristics at various steady-state and unstable operating conditions. Flow phenomenology is discussed throughout the various operating regimes and data and image analysis techniques are introduced. Combustion chamber operation very near to the critical point has highlighted very stable operating conditions in terms of propellant flow, flame stability and combustor response.