Publication Date:
2013-08-31
Description:
Nonintrusive measurements were made of a normal shock wave/boundary layer interaction. Two dimensional measurements were made throughout the interaction region while 3-D measurements were made in the vicinity of the shock wave. The measurements were made in the corner of the test section of a continuous supersonic wind tunnel in which a normal shock wave had been stabilized. Laser Doppler Anemometry, surface pressure measurement and flow visualization techniques were employed for two freestream Mach number test cases: 1.6 and 1.3. The former contained separated flow regions and a system of shock waves. The latter was found to be far less complicated. The results define the flow field structure in detail for each case.
Keywords:
AERODYNAMICS
Type:
NASA, Langley Research Center, Transonic Symposium: Theory, Application, and Experiment, Volume 1, Part 2; p 741-764
Format:
application/pdf
