XFOIL Version 6.94 Calculated polar for: FX 66-182 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 0.4292 0.01527 0.00823 -0.0856 0.5634 0.6635 0.500 0.4841 0.01539 0.00833 -0.0859 0.5590 0.6698 1.000 0.5404 0.01556 0.00841 -0.0865 0.5550 0.6758 1.500 0.5967 0.01563 0.00844 -0.0871 0.5513 0.6805 2.000 0.6521 0.01616 0.00894 -0.0876 0.5470 0.6860 2.500 0.7044 0.01628 0.00916 -0.0875 0.5430 0.6917 3.000 0.7573 0.01656 0.00949 -0.0876 0.5385 0.6969 3.500 0.8107 0.01667 0.00968 -0.0877 0.5339 0.7026 4.000 0.8664 0.01686 0.00991 -0.0882 0.5300 0.7083 4.500 0.9226 0.01744 0.01048 -0.0890 0.5261 0.7143 5.000 0.9705 0.01789 0.01113 -0.0883 0.5222 0.7207 5.500 1.0180 0.01824 0.01169 -0.0875 0.5170 0.7277 6.000 1.0686 0.01857 0.01216 -0.0872 0.5123 0.7352 6.500 1.1227 0.01885 0.01257 -0.0875 0.5085 0.7430 7.000 1.1783 0.01928 0.01309 -0.0880 0.5034 0.7522 7.500 1.2174 0.01933 0.01344 -0.0856 0.4956 0.7637 8.000 1.2746 0.01853 0.01273 -0.0857 0.4870 0.7777 8.500 1.3182 0.01833 0.01278 -0.0838 0.4778 0.7957 9.000 1.3646 0.01710 0.01168 -0.0816 0.4630 0.8277 10.000 1.4101 0.01665 0.01169 -0.0703 0.4243 1.0000 10.500 1.4175 0.01738 0.01246 -0.0629 0.4010 1.0000 11.000 1.4090 0.01915 0.01422 -0.0545 0.3660 1.0000 11.500 1.3768 0.02305 0.01795 -0.0458 0.3207 1.0000 12.000 1.3282 0.02965 0.02435 -0.0390 0.2689 1.0000 12.500 1.2747 0.03791 0.03241 -0.0341 0.2230 1.0000 13.000 1.2276 0.04659 0.04089 -0.0310 0.1769 1.0000 13.500 1.1847 0.05575 0.04976 -0.0292 0.1244 1.0000 14.000 1.1493 0.06487 0.05854 -0.0285 0.0716 1.0000 14.500 1.1287 0.07295 0.06639 -0.0284 0.0397 1.0000 15.000 1.1242 0.07959 0.07306 -0.0287 0.0312 1.0000 15.500 1.1236 0.08601 0.07957 -0.0294 0.0274 1.0000 16.000 1.1226 0.09267 0.08632 -0.0304 0.0252 1.0000 16.500 1.1292 0.09840 0.09220 -0.0314 0.0237 1.0000 17.000 1.1345 0.10435 0.09822 -0.0327 0.0222 1.0000 17.500 1.1447 0.10957 0.10356 -0.0339 0.0208 1.0000 18.000 1.1572 0.11441 0.10849 -0.0350 0.0196 1.0000 18.500 1.1761 0.11778 0.11182 -0.0355 0.0186 1.0000 19.000 1.1895 0.12262 0.11688 -0.0369 0.0174 1.0000 19.500 1.2059 0.12668 0.12099 -0.0382 0.0162 1.0000 20.000 1.2156 0.13193 0.12644 -0.0402 0.0147 1.0000 20.500 1.2306 0.13600 0.13052 -0.0418 0.0134 1.0000 21.000 1.2337 0.14240 0.13725 -0.0445 0.0125 1.0000 21.500 1.2378 0.14861 0.14374 -0.0474 0.0118 1.0000 22.000 1.2440 0.15416 0.14942 -0.0502 0.0112 1.0000