Study of a Supersonic Combustor Employing an Aerodynamic Ramp Pilot Injector

Calculations of the fuel-air mixing characteristics of a scramjet combustor concept employing an aerodynamic ramp fuel injector coupled with a cavity flameholder have been performed. The calculations of the nonreacting flow through the combustor represent a first step in a collab- orative computational and experimental effort to investi- gate the performance of this scramjet combustor concept. The nonreacting flow through the combustor yielded a high total pressure recovery. Entrainment of fuel within the cavity produced a large volume of the cavity that had a local fuel-air equivalence ratio between 0.5 and 2.5, which did not vary appreciably as the overall fuel- air equivalence ratio increased from 0.13 to 0.72. How- ever, the stronger shock and expansion waves generated at the higher overall fuel-air equivalence ratios caused the fuel distribution within the cavity to become more three- dimensional. A shock train upstream of the fuel injectors was generated by imposing a back pressure at the exit plane of the domain. The shock train separated the side wall boundary layer and led to a dramatic increase in the fuel-air mixing. The mixing was characterized by en- hanced asymmetric spreading toward the side wall and was accompanied by an increased loss in total pressure through the combustor.