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Abstract <p>The ignition process in supersonic flows is complex, influenced by the fuel ratios, design, and flow conditions. The cavity combustors stabilize regions for ignition, whereas ramps aid in mixing fuel and air. The shock waves generated by these ramps interact with the fuel stream, enhancing micro-mixing and creating recirculation zones to improve fuel-air mixing. In supersonic flow, these vortices contribute to macro-mixing, while the interaction between shocks and the fuel stream generates boroclinic torque at the air-fuel interface, further improving micromixing. A full-scale experimental combustor facility is designed and tests are done on a supersonic combustor incorporated with ramps and cavities, using aviation kerosene. The combustor wall pressures and temperatures are obtained using the instrumentation. Numerical studies examined the effects of the Mach number on flow. ANSYS-based simulations revealed that the combustor configuration and the Mach number play a vital role. It is inferred that a staged injection improves mixing and thrust. At a combustor entry Mach number equal to 2, in the diverging section the static pressures initially rise, then decrease with time, indicating supersonic combustion, with a wall pressure increase of about 1.3 bar the exit static pressures rise to equalize with ambient conditions. The adopted strategies resulted in sustained supersonic combustion.</p>

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Investigations on a Supersonic Combustor with Ramps and Cavities for Sustained Combustion

  • 期刊论文

  • 发表日期：

    2025年2月1日

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  • G. Amba Prasad Rao,
  • J.V.S. Moorthy

摘要

Abstract

The ignition process in supersonic flows is complex, influenced by the fuel ratios, design, and flow conditions. The cavity combustors stabilize regions for ignition, whereas ramps aid in mixing fuel and air. The shock waves generated by these ramps interact with the fuel stream, enhancing micro-mixing and creating recirculation zones to improve fuel-air mixing. In supersonic flow, these vortices contribute to macro-mixing, while the interaction between shocks and the fuel stream generates boroclinic torque at the air-fuel interface, further improving micromixing. A full-scale experimental combustor facility is designed and tests are done on a supersonic combustor incorporated with ramps and cavities, using aviation kerosene. The combustor wall pressures and temperatures are obtained using the instrumentation. Numerical studies examined the effects of the Mach number on flow. ANSYS-based simulations revealed that the combustor configuration and the Mach number play a vital role. It is inferred that a staged injection improves mixing and thrust. At a combustor entry Mach number equal to 2, in the diverging section the static pressures initially rise, then decrease with time, indicating supersonic combustion, with a wall pressure increase of about 1.3 bar the exit static pressures rise to equalize with ambient conditions. The adopted strategies resulted in sustained supersonic combustion.