Circular Scramjet Combustor Transition Lengths at Mach 8 and 12.8 Flight Conditions

This paper presents an experimental study of laminar-to-turbulent boundary-layer transition in a circular duct at high supersonic Mach number. The experiments were conducted as an essential precursor for scramjet drag reduction studies using a direct-connect scramjet combustor. The experiments were carried out at the University of Queensland T4 shock tunnel using a 33.2-mm-diam and 500-mm-long circular duct at flow stagnation enthalpies of 3.2 and 7.8 MJ/k g. The experimental work was supported with two-dimensional and three-dimensional Reynolds-averaged Navier-Stokes simulations. Comparisons between numerical and experimental results show initial rise in heat transfer at the first shock-wave boundary-layer interaction, followed by an unsteady laminar region interspersed with possible turbulent spots. Heat transfer measurements along the wall that show sustained transition to turbulence did not occur within the full length of the duct. Transition to turbulence within the circular duct is significantly delayed when compared with estimations based on flat-plate results. These results indicate that there is value in revisiting the assumption of a fully turbulent boundary layer for previous semi-direct-connect scramjet combustion experiments that have used transition length correlations derived from flat-plate experiments within the same facility. The findings from this investigation have implications for conclusions drawn from experimental investigations of boundary-layer-combustion-induced drag reduction in circular ducts.