Supersonic Combustion Tests with a Double-Oblique-Shock Scramjet in a Shock Tunnel

摘要

Some results of a continuing research program to develop a capability for testing integrated scramjets are reported. During this research program, an integrated double-oblique-shock scramjet model was developed to provide a test bed for supersonic combustion tests and for instrumentation development essential for analysis of combustion test results. Results are presented for tests in which hydrogen fuel was injected into the combustor. Injection of the fuel, from sonic orifices in the wall, normal to the flow did not lead to satisfactory combustion data, supposedly because of the cold boundary layer. Injection through sonic orifices in a series of diamond airfoil injectors led to combustion confirmed by all the following measurements: (1) an increase in static pressure within the combustor downstream of the injection station, (2) an increase in surface heat-transfer rate, (3) an increase in static temperature as measured by the sodium line reversal technique, (4) an increase in output of radiation sensor gages, and (5) a decrease in flow Mach number inferred from static to pitot pressure measurements. The measured increases were proportionate to increases in computed average equivalence ratio. The combustion results are compared with numerical solutions. In general, the measured temperatures and pressures in the combustor with heat addition were higher than the calculated values. (Author)