Experimental Study on Effect of Boundary Layer on Combustion Modes in a Supersonic Combustor

Combustion tests were conducted to reveal the mechanism and the dominant factor for attainment of the dual-mode operation in a supersonic combustor in a directly-connected wind tunnel facility which supplied Mach 2.5 airflow with a total temperature of 1500 K and a total pressure of 1.0 MPa. In order to extract the dominant factor, boundary layer bleeding technique was applied to the supersonic combustor, and the combustion mode (i.e., supersonic combustion or dual-mode combustion) was drastically and actively changed. In the case of dual-mode without bleeding, the combustion mode was transited to supersonic combustion by the bleeding, and a further increase in fuel flow rate with bleeding resulted in little or no change in the wall pressure distribution. On the other hand, in the case of supersonic combustion mode without bleeding, the bleeding resulted in lower or higher wall pressure, while the separation point did not vary so much; the reason was still uncertain.