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 = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3361 0.01066 0.00450 -0.0761 0.6894 0.6798 0.500 0.3933 0.01067 0.00453 -0.0764 0.6770 0.6929 1.000 0.4506 0.01082 0.00466 -0.0766 0.6658 0.7066 1.500 0.5069 0.01085 0.00475 -0.0767 0.6535 0.7192 2.000 0.5642 0.01102 0.00492 -0.0770 0.6418 0.7342 2.500 0.6197 0.01109 0.00507 -0.0767 0.6292 0.7475 3.000 0.6754 0.01125 0.00531 -0.0767 0.6171 0.7620 3.500 0.7317 0.01139 0.00548 -0.0767 0.6041 0.7778 4.000 0.7845 0.01145 0.00569 -0.0759 0.5875 0.7926 4.500 0.8373 0.01156 0.00590 -0.0752 0.5702 0.8084 5.000 0.8892 0.01163 0.00605 -0.0742 0.5494 0.8251 5.500 0.9378 0.01168 0.00615 -0.0726 0.5195 0.8435 6.000 0.9848 0.01188 0.00639 -0.0708 0.4912 0.8624 6.500 1.0272 0.01209 0.00674 -0.0681 0.4574 0.8837 7.000 1.0600 0.01249 0.00712 -0.0636 0.4086 0.9096 7.500 1.0772 0.01322 0.00764 -0.0565 0.3309 0.9571 8.000 1.0983 0.01553 0.00926 -0.0531 0.2103 1.0000 8.500 1.1014 0.01814 0.01130 -0.0467 0.1231 1.0000 9.000 1.1052 0.02092 0.01372 -0.0411 0.0773 1.0000 9.500 1.1174 0.02350 0.01623 -0.0372 0.0605 1.0000 10.000 1.1317 0.02619 0.01896 -0.0340 0.0521 1.0000 10.500 1.1445 0.02923 0.02205 -0.0312 0.0463 1.0000 11.000 1.1559 0.03261 0.02551 -0.0288 0.0418 1.0000 11.500 1.1662 0.03632 0.02927 -0.0268 0.0379 1.0000 12.000 1.1783 0.04009 0.03308 -0.0253 0.0346 1.0000 12.500 1.1939 0.04380 0.03694 -0.0239 0.0318 1.0000 13.000 1.2128 0.04744 0.04056 -0.0228 0.0296 1.0000 13.500 1.2313 0.05122 0.04457 -0.0221 0.0277 1.0000 14.000 1.2551 0.05471 0.04807 -0.0215 0.0262 1.0000 14.500 1.2770 0.05881 0.05242 -0.0209 0.0252 1.0000 15.000 1.2912 0.06379 0.05773 -0.0209 0.0244 1.0000 15.500 1.3011 0.06931 0.06352 -0.0214 0.0236 1.0000 16.000 1.3081 0.07546 0.06994 -0.0224 0.0231 1.0000 16.500 1.3139 0.08185 0.07652 -0.0240 0.0225 1.0000 17.000 1.3109 0.09021 0.08517 -0.0262 0.0221 1.0000 17.500 1.2842 0.10169 0.09712 -0.0313 0.0220 1.0000 18.000 1.2553 0.11458 0.11043 -0.0380 0.0220 1.0000 18.500 1.2252 0.12857 0.12480 -0.0460 0.0221 1.0000