An experimental study of supersonic combustion with incoming high temperature pure air stream obtained by shock tunnel
An experimental study is conducted to investigate the phenomena of supersonic combustion by means of shock tunnel. By using shock tunnel, test air is compressed by reflected shock wave up to stagnation temperature of 2800 K and stagnation pressure of 0.35 MPa. Heated air is used as a reservoir gas of supersonic nozzle. Hydrogen is injected transversely through circular hole into freestream of Mach 2. Flow duration is 430 microseconds. By schlieren method and CCD UV camera, the effects of injection pressure to flowfield and the combustion were investigated.
KeywordsInjection Pressure Supersonic Flow Circular Hole Stagnation Pressure Flow Duration
Unable to display preview. Download preview PDF.
- 2.R.C. Roger, D.P. Capriotti and R.W. Guy: Experimental supersonic combustion research at NASA Langley. AIAA-Paper 98-2506 (1998)Google Scholar
- 3.G. Smeets, C. Quenett: ‘Shock Tube Investigation of H2 Combustion in a High Temperature Supersonic Air Flow’. In: (Scramjet), Proceeding of IUTAM Symposium on Combustion in Supersonic Flows, Netherlands, 1997, pp. 173–178Google Scholar