Experimental Investigations of Scramjet Combustor Characteristics

Abstract

The design of scramjet engines requires systematic experimental data. Hence, in the present study, a scramjet combustor model was investigated experimentally using the facil- ities at the Institute of Aerospace Thermodynamics, UniversitStuttgart, Germany. The facilities can deliver a continuous supersonic flow of Mach number 2.1 with a maximum total temperature of 1500K and a maximum total pressure of 1MPa. This corresponds to a flight Mach number of 5 at an altitude of about 30km. In the present experiments, hydrogen was injected using a central injector concept that enhances mixing efficiency by inducing streamwise vortices. Parameters used in this study include injector type, injection position, combustor opening angle, total temperature and equivalence ratio. Measurements of static wall pressure distributions and high speed pictures of the flow field indicated the presence of two distinctive combustion modes. These were dependent on running condi- tions. Weak combustion, which is accompanied by small pressure rise, showed a detached flame. Strong combustion, where the flame is attached to the injector, showed a higher pressure rise. The transition between the two modes was very sensitive to the test con- ditions at the investigated Mach number. The tansition mechanism was supposed to be a detonation wave travelling upstream to the trailing edge of the injector. After mode transition occured, a stable combustion was achieved, where the central injector acted as a kind of flameholder. Due to the geometric blockage of the injector and the growing bound- ary layer the static pressure in the isolator increased. In all test cases this pressure rise causes a pre-combustion shock train structure. In the present study the central injection concept was found to have good mixing and combustion capabilities. Especially if strong combustion is active and combustor height is sufficient to reduce geometric blockage effects.