XFOIL Version 6.94 Calculated polar for: NACA 63-412 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3427 0.00859 0.00291 -0.0775 0.6605 0.6494 0.500 0.4008 0.00870 0.00294 -0.0779 0.6490 0.6623 1.000 0.4582 0.00874 0.00306 -0.0782 0.6372 0.6752 1.500 0.5162 0.00889 0.00313 -0.0786 0.6252 0.6882 2.000 0.5730 0.00893 0.00330 -0.0787 0.6131 0.7015 2.500 0.6302 0.00911 0.00345 -0.0790 0.6009 0.7152 3.000 0.6868 0.00919 0.00365 -0.0791 0.5886 0.7290 3.500 0.7425 0.00932 0.00384 -0.0791 0.5736 0.7430 4.000 0.7980 0.00949 0.00400 -0.0789 0.5562 0.7576 4.500 0.8518 0.00961 0.00421 -0.0785 0.5359 0.7726 5.000 0.9042 0.00974 0.00443 -0.0778 0.5090 0.7880 5.500 0.9561 0.00998 0.00472 -0.0771 0.4800 0.8039 6.000 1.0059 0.01036 0.00510 -0.0760 0.4474 0.8214 6.500 1.0506 0.01092 0.00559 -0.0741 0.3925 0.8392 7.000 1.0814 0.01222 0.00649 -0.0701 0.2943 0.8601 7.500 1.0971 0.01410 0.00785 -0.0639 0.1863 0.8853 8.000 1.1015 0.01555 0.00905 -0.0553 0.1200 0.9273 8.500 1.1134 0.01725 0.01050 -0.0493 0.0731 1.0000 9.000 1.1328 0.01912 0.01225 -0.0452 0.0539 1.0000 9.500 1.1532 0.02104 0.01417 -0.0417 0.0454 1.0000 10.000 1.1719 0.02322 0.01639 -0.0384 0.0401 1.0000 11.000 1.2034 0.02856 0.02189 -0.0327 0.0321 1.0000 11.500 1.2151 0.03186 0.02528 -0.0304 0.0288 1.0000 12.000 1.2252 0.03552 0.02899 -0.0285 0.0260 1.0000 12.500 1.2411 0.03890 0.03252 -0.0272 0.0236 1.0000 13.000 1.2497 0.04315 0.03684 -0.0262 0.0218 1.0000 13.500 1.2627 0.04725 0.04112 -0.0256 0.0205 1.0000 14.000 1.2743 0.05171 0.04568 -0.0255 0.0193 1.0000 14.500 1.2804 0.05702 0.05110 -0.0256 0.0185 1.0000 15.000 1.2906 0.06213 0.05642 -0.0260 0.0179 1.0000 15.500 1.2994 0.06767 0.06215 -0.0270 0.0172 1.0000 16.000 1.3066 0.07362 0.06828 -0.0282 0.0168 1.0000 16.500 1.3127 0.07998 0.07481 -0.0301 0.0163 1.0000 17.000 1.3173 0.08675 0.08173 -0.0322 0.0160 1.0000 17.500 1.3201 0.09390 0.08903 -0.0346 0.0156 1.0000 18.000 1.3177 0.10228 0.09769 -0.0380 0.0154 1.0000 18.500 1.3105 0.11181 0.10754 -0.0426 0.0152 1.0000 19.000 1.2982 0.12265 0.11870 -0.0484 0.0150 1.0000 19.500 1.2807 0.13479 0.13118 -0.0555 0.0149 1.0000 20.000 1.2559 0.14897 0.14572 -0.0645 0.0149 1.0000 20.500 1.2211 0.16615 0.16327 -0.0761 0.0149 1.0000 21.000 1.1541 0.19329 0.19086 -0.0948 0.0153 1.0000