A sidewall-compression-type scramjet engine was tested under Mach 4 flight conditions. The tested engine had an inlet, a constant cross-sectional area isolator, a constant cross-sectional area combustor, a diverging combustor, and an internal nozzle. In a previous study under the same flight conditions, the maximum thrust increment using fuel injection within the constant-area combustor was 1,380 N at an equivalence ratio of 0.31, and further fuel injection resulted in combustor-inlet interaction (designated as CII). To suppress the CII in the present study, we attempted (1) two-stage fuel injection within the constant-area combustor and the diverging combustor and (2) a boundary layer bleed on the top wall. The former was to suppress heat release around the first-stage fuel injectors in the constant-area combustor, and the latter was to decrease interaction length by decreasing the boundary layer thickness on the top wall. In the case of two-stage fuel injection, the maximum thrust increment was 2,230 N at an equivalence ratio of 0.63. In the case of the boundary layer bleed, on the other hand, the maximum thrust increment was 2,300 N at an equivalence ratio of 0.66. Thus, two-stage fuel injection and boundary layer bleed led to 62% and 67% higher maximum thrust increments than that obtained in the previous study, respectively. Finally, both the two-stage fuel injection and boundary layer bleed were applied simultaneously to obtain the best thrust performance, and the maximum thrust increment was 2,560 N at an equivalence ratio of 0.95. As a result, we obtained an 86% higher maximum thrust increment than that in the previous study. The thrust achievement factor, which was defined as the ratio of the maximum thrusts obtained from experiment and theoretical prediction, under this condition was estimated as 70%.
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