XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0301 0.05549 0.02259 -0.0213 0.9994 1.0006 -2.750 -0.0098 0.05558 0.02230 -0.0209 0.9994 1.0006 -2.500 0.0104 0.05575 0.02209 -0.0205 0.9994 1.0006 -2.250 0.0304 0.05598 0.02202 -0.0201 0.9994 1.0006 -2.000 0.0502 0.05627 0.02205 -0.0197 0.9994 1.0006 -1.750 0.0697 0.05662 0.02218 -0.0192 0.9994 1.0006 -1.500 0.0891 0.05704 0.02238 -0.0187 0.9994 1.0006 -1.250 0.1084 0.05753 0.02271 -0.0183 0.9994 1.0006 -1.000 0.1274 0.05807 0.02313 -0.0179 0.9994 1.0006 -0.750 0.1462 0.05868 0.02365 -0.0174 0.9994 1.0006 -0.500 0.1648 0.05937 0.02425 -0.0170 0.9994 1.0006 -0.250 0.1831 0.06012 0.02496 -0.0166 0.9994 1.0006 0.000 0.2012 0.06095 0.02579 -0.0163 0.9994 1.0006 0.250 0.2190 0.06186 0.02672 -0.0160 0.9994 1.0006 0.500 0.2366 0.06286 0.02776 -0.0158 0.9994 1.0006 0.750 0.2538 0.06396 0.02893 -0.0156 0.9994 1.0006 1.000 0.2706 0.06515 0.03024 -0.0155 0.9994 1.0006 1.250 0.2869 0.06646 0.03168 -0.0156 0.9994 1.0006 1.500 0.3025 0.06789 0.03327 -0.0157 0.9994 1.0006 1.750 0.3174 0.06947 0.03503 -0.0159 0.9994 1.0006 2.000 0.3313 0.07120 0.03697 -0.0162 0.9994 1.0006 2.250 0.3442 0.07313 0.03912 -0.0167 0.9994 1.0006 2.500 0.3559 0.07528 0.04150 -0.0174 0.9994 1.0006 2.750 0.3662 0.07768 0.04413 -0.0183 0.9994 1.0006 3.000 0.3751 0.08032 0.04698 -0.0194 0.9994 1.0006 3.250 0.3827 0.08321 0.05005 -0.0207 0.9994 1.0006 3.500 0.3894 0.08628 0.05328 -0.0222 0.9994 1.0006 3.750 0.3957 0.08948 0.05661 -0.0238 0.9994 1.0006 4.000 0.4019 0.09275 0.05999 -0.0255 0.9994 1.0006 4.250 0.4085 0.09607 0.06342 -0.0272 0.9994 1.0006 4.500 0.4154 0.09942 0.06686 -0.0290 0.9994 1.0006 4.750 0.4226 0.10276 0.07034 -0.0308 0.9994 1.0006 5.000 0.4301 0.10610 0.07376 -0.0327 0.9994 1.0006