Experimental Investigation on Performance of an Arc Transverse Injection in a Supersonic Combustor Flow

The design of a supersonic combustor ramjet engine has gained significant attention for futuristic air-breathing engines. Achieving efficient fuel mixing and complete combustion within a short period poses a substantial challenge. This study has directed efforts towards improving mixing and minimizing total pressure losses across the supersonic combustor to enhance performance. Experiments were conducted to investigate the influence of fuel injection geometry in the supersonic combustor with an entry Mach number of 2. Two different fuel orifice geometries, inclined at 45°, were considered. The investigation covered four different momentum flux ratios: 0.8, 1.0, 1.2, and 1.4 respectively. Various measurements were conducted to observe flow phenomena inside the supersonic combustion, including wall static pressure measurement, Schlieren visualization, exit Mach number and total pressure loss measurement. Without injection cases exhibited weaker compression and expansion inside the combustor. During injection, the rise in wall pressure indicated that the bow shock formed in front of the arc injection was slightly weaker than that of the circular injection. The impinging bow shock on the opposite wall also exhibited higher strength, resulting in a static pressure rise. As a result, the lower total pressure ratio across the shock indicates a higher momentum exchange between the main flow and the orifice. Therefore, arc injection has proven more effective in exchanging momentum inside the supersonic combustor. Consequently, the Mach number at the exit of the combustor was higher for the arc injection.