XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY MS(1)-0317 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.500 -0.8667 0.10357 0.09843 -0.0443 1.0000 0.0279 -17.250 -0.8970 0.09361 0.08822 -0.0504 1.0000 0.0280 -17.000 -0.9154 0.08637 0.08079 -0.0547 1.0000 0.0281 -16.750 -0.9287 0.08034 0.07460 -0.0581 1.0000 0.0283 -16.500 -0.9389 0.07508 0.06918 -0.0609 1.0000 0.0285 -16.250 -0.9470 0.07038 0.06433 -0.0633 1.0000 0.0288 -16.000 -0.9534 0.06615 0.05996 -0.0652 1.0000 0.0291 -15.750 -0.9583 0.06230 0.05595 -0.0667 1.0000 0.0294 -15.500 -0.9619 0.05879 0.05227 -0.0678 1.0000 0.0298 -15.250 -0.9644 0.05558 0.04889 -0.0686 1.0000 0.0303 -15.000 -0.9657 0.05263 0.04574 -0.0692 1.0000 0.0307 -14.750 -0.9639 0.05012 0.04309 -0.0694 1.0000 0.0311 -14.500 -0.9580 0.04812 0.04108 -0.0694 1.0000 0.0315 -14.250 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