Combustion characteristics in a supersonic combustor with ethylene injection upstream of dual parallel cavities were investigated experimentally in a direct-connected test rig with inflow conditions of Ma = 3.46, Po = 3.60 Mpa and To = 1430 K. The combustor has a two-dimensional rectangular configuration with single-side expansion. Two open cavities with the same size were mounted on the expanded and horizontal walls, respectively. Static pressure distribution in the axial direction was measured along the centerline of the expanded wall. High-speed flame luminosity and schlieren were used to capture the combustion and flow structures at different equivalence ratios. Two clusters of separated and asymmetric flames were found to be stabilized near the dual parallel cavities in all tests. The flame and flow characteristics changed with the combustion modes. For scramjet mode, no obvious flow separation occurred near the walls and the two flames were both stabilized in the cavity shear layer and recirculation zone. For ramjet mode, the high back pressure resulting from intense heat release induced a large-scale recirculation zone upstream of the cavity mounted on the expanded wall, which supplied a favorable combustion condition and the flame was stabilized in the jet-wake. Meanwhile, there was no obvious separation near the horizontal wall, with the local flame stabilized in the cavity shear layer. It is suggested that the combustion near the horizontal wall should be enhanced to improve the combustion performance and avoid a non-uniform flow field at the combustor exit.

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Dynamic combustion characteristics of a rectangular scramjet combustor with single-side expansion were studied experimentally and numerically. Experiments were implemented with an isolator entrance Mach number of 3.46, and an air stagnation temperature of 1430 K. Ethylene was utilized to fuel the combustor over an equivalence ratio range of 0.20 < φ < 0.63. Results indicated that the combustion modes varied from different equivalence ratios. For an intermediate φ = 0.375, an intermittent dynamic combustion occurred. During the dynamic process, the flame sometimes stabilized in the jet wake of the top cavity, and at other time it oscillated between dual parallel cavities. The pseudo-shock train traveled periodically along the length of the combustor, and the penetration depths of the two injectors exchanged. Quantitative analysis illustrated that the average frequency of unsteady combustion was approximately 200 Hz. The reason for the occurrence of the self-sustained dynamic process was related to the interactions between the shock-induced separated region and heat release.

Supersonic combustion with distributed injection of supercritical kerosene in a model scramjet engine was experimentally investigated in a Mach 2.92 facility with stagnation temperatures of approximately 1430 K. Multicavities were used to stabilize and enhance the combustion in the supersonic combustor. Supercritical kerosene at temperatures of approximately 780 K was prepared using a heat exchanger driven by the hot gas from a preburner and injected into the combustor at equivalence ratios of 1.0. Static pressure distribution in the axial direction was measured along the centerline of the model combustor top walls. A high-speed imaging camera was used to capture flame luminosity and combustion region distribution. In the experiments, combustor performances with different injection locations, injection stages, cavity locations, and numbers of cavities were investigated systematically. The experimental results showed that the injection penetration and local combustion had a strong coupling with the upstream flow. The combustion region and heat release distribution changes obviously due to the various cavityinjection schemes, and the combustion performance could be improved when the injection location and distribution of supercritical kerosene mass flow rate were optimized. Nomenclature A= cavity rear wall angle Bn = nth cavity on bottom wall of the scramjet combustor Bnf = injection set upstream of the cavity Bn front wall D = cavity depth L = cavity length LD = cavity length-to-depth ratio Ma = Mach number P 0i = stagnation pressure of the supercritical kerosene Tn = nth cavity on top wall of the scramjet combustor Tnf = injection set upstream of the cavity Tn front wall Tnr = injection set downstream of the cavity Tn rear wall T 0i = stagnation temperature of the supercritical kerosene ψ = equivalence ratio