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Supersonic vortex breakdown over a delta wing in transonic flowThe effects of freestream Mach number and angle of attack on the leading-edge vortex breakdown due to the terminating shock on a 65-degree, sharp-edged, cropped delta wing are investigated computationally, using the time-accurate solution of the laminar unsteady compressible full Navier-Stokes equations with the implicit upwind flux-difference splitting, finite-volume scheme. A fine O-H grid consisting of 125 x 85 x 84 points in the wrap-around, normal, and axial directions, respectively, is used for all the flow cases. Keeping the Reynolds number fixed at 3.23 x 10 exp 6, the Mach number is varied from 0.85 to 0.9 and the angle of attack is varied from 20 to 24 deg. The results show that, at 20-deg angle of attack, the increase of the Mach number from 0.85 to 0.9 results in moving the location of the terminating shock downstream. The results also show that, at 0.85 Mach number, the increase of the angle of attack from 20 to 24 deg results in moving the location of the terminating shock upstream. The results are in good agreement with the experimental data.
Document ID
19930063254
Document Type
Conference Paper
Authors
Kandil, Hamdy A. (NASA Langley Research Center Hampton, VA, United States)
Kandil, Osama A. (Old Dominion Univ. Norfolk, VA, United States)
Liu, C. H. (NASA Langley Research Center Hampton, VA, United States)
Date Acquired
August 16, 2013
Publication Date
January 1, 1993
Publication Information
Publication: In: AIAA Applied Aerodynamics Conference, 11th, Monterey, CA, Aug. 9-11, 1993, Technical Papers. Pt. 2 (A93-47201 19-02)
Subject Category
AERODYNAMICS
Report/Patent Number
AIAA PAPER 93-3472
Funding Number(s)
CONTRACT_GRANT: NAG1-994
Distribution Limits
Public
Copyright
Other