The range of applicability of classical panel methods for incompressible flows has often been extended to address compressible, subsonic flows by "compressibility corrections", where the solution for an incompressible flow about a body (or a modified shape of body) is used to predict the compressible results about the original shape - the classical approach is to use the Prandtl-Glauert/Göthert approach. However, this has met with mixed success and requires care in its correct application. This has led other methods to use corrections to the incompressible solution for the pressures about the body (e.g. the von Kármán-Tsien compressibility correction).

Recent work at Flow Solutions has obtained a new formulation of the compressibility correction that yields accurate answers up to the onset of transonic flow and does not suffer from the problems seen with other approaches.

Presented here is a comparison of NEWPAN with experimental results, for a wing plus fuselage test case in subsonic compressible flow. This testcase was reported by W.Schneider at DLR [1], and features a wing of Aspect Ratio 6 with a 9% thick RAE101 profile. Experimentally it was tested at Mach numbers from 0.5 to 0.8.

NEWPAN results with Prandtl-Glauert compressibility correction show excellent agreement with experiment - aside from the highest Mach number/incidence combinations, where the experimental data shows that a shock has started to form. Cpcrit, the value of pressure coefficient necessary to achieve local sonic flow, is plotted for the high Mach number cases. Where the sectional pressure distribution crosses the Cpcrit line we would expect poorer experimental comparisons due to the presence of locally sonic flow.

[1] W.Schneider, Druckverteilungsmessungen an Pfeilflügel-Rumpf-Anordnungen bei hohen Unterschall-Machzahlen, Teil II: Flügel mit dickem Rumpf, DFVLR, Aerodynamische Versuchsanstalt Göttingen, interner Bericht 70 A 40, 1970

Use the interactive NEWPAN results browser to view the comparison with experimental data.

alpha = 0 alpha = 2 alpha = 4 alpha = 6 Mach=0.0 Mach=0.5 Mach=0.6 Mach=0.7 Mach=0.8

To change the spanwise section displayed in the graph plot, left mouse click on a white wing section in the 3D view above.