XFOIL Version 6.96 Calculated polar for: FX 61-163 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.250 -0.6236 0.05884 0.05593 -0.0661 1.0001 0.0052 -12.000 -0.6492 0.05279 0.04966 -0.0681 1.0001 0.0052 -11.750 -0.6689 0.04815 0.04483 -0.0688 1.0001 0.0052 -11.500 -0.6846 0.04439 0.04089 -0.0685 1.0001 0.0052 -11.250 -0.6923 0.04076 0.03705 -0.0690 0.9969 0.0052 -11.000 -0.6919 0.03746 0.03354 -0.0701 0.9764 0.0053 -10.750 -0.6496 0.03302 0.02870 -0.0806 0.9289 0.0054 -10.500 -0.6120 0.03007 0.02521 -0.0874 0.8522 0.0055 -10.250 -0.6019 0.02856 0.02333 -0.0867 0.8057 0.0056 -10.000 -0.5903 0.02717 0.02164 -0.0860 0.7766 0.0058 -9.750 -0.5769 0.02590 0.02012 -0.0852 0.7546 0.0059 -9.500 -0.5599 0.02469 0.01867 -0.0849 0.7374 0.0060 -9.250 -0.5413 0.02355 0.01732 -0.0845 0.7234 0.0062 -9.000 -0.5216 0.02249 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1.1827 0.01200 0.00563 -0.1101 0.3649 0.6947 6.750 1.2318 0.01267 0.00626 -0.1095 0.3335 0.6984 7.000 1.2536 0.01319 0.00667 -0.1088 0.3063 0.7002 7.250 1.2730 0.01386 0.00719 -0.1077 0.2716 0.7019 7.500 1.2900 0.01466 0.00780 -0.1063 0.2345 0.7034 7.750 1.3056 0.01553 0.00849 -0.1047 0.1964 0.7048 8.000 1.3199 0.01643 0.00924 -0.1029 0.1652 0.7063 8.250 1.3329 0.01737 0.01002 -0.1010 0.1353 0.7079 8.500 1.3447 0.01830 0.01083 -0.0988 0.1098 0.7098 8.750 1.3546 0.01914 0.01159 -0.0963 0.0908 0.7117 9.000 1.3637 0.02003 0.01242 -0.0937 0.0742 0.7136 9.500 1.3811 0.02205 0.01437 -0.0891 0.0513 0.7167 9.750 1.3901 0.02309 0.01545 -0.0871 0.0420 0.7182 10.250 1.3976 0.02618 0.01851 -0.0828 0.0232 0.7217 10.500 1.4057 0.02762 0.02002 -0.0816 0.0199 0.7237 10.750 1.4126 0.02929 0.02173 -0.0805 0.0172 0.7256 11.000 1.4209 0.03092 0.02344 -0.0797 0.0153 0.7274 11.250 1.4269 0.03284 0.02542 -0.0789 0.0128 0.7291 11.500 1.4321 0.03490 0.02753 -0.0782 0.0100 0.7308 11.750 1.4357 0.03714 0.02983 -0.0776 0.0070 0.7323 12.000 1.4372 0.03964 0.03240 -0.0769 0.0047 0.7339 12.250 1.4393 0.04209 0.03493 -0.0764 0.0040 0.7358 12.500 1.4413 0.04458 0.03752 -0.0758 0.0036 0.7378 12.750 1.4426 0.04717 0.04019 -0.0753 0.0033 0.7398 13.000 1.4433 0.04982 0.04294 -0.0749 0.0031 0.7418 13.250 1.4431 0.05258 0.04580 -0.0744 0.0029 0.7436 13.500 1.4428 0.05544 0.04875 -0.0741 0.0028 0.7454 13.750 1.4431 0.05825 0.05167 -0.0738 0.0027 0.7471 14.000 1.4432 0.06117 0.05470 -0.0737 0.0027 0.7488 14.250 1.4425 0.06426 0.05791 -0.0736 0.0026 0.7508 14.500 1.4410 0.06750 0.06126 -0.0737 0.0025 0.7528 14.750 1.4399 0.07075 0.06462 -0.0738 0.0025 0.7549 15.000 1.4379 0.07414 0.06813 -0.0740 0.0025 0.7570 15.250 1.4352 0.07773 0.07185 -0.0743 0.0024 0.7593 15.500 1.4318 0.08148 0.07571 -0.0748 0.0024 0.7615 15.750 1.4280 0.08533 0.07968 -0.0753 0.0023 0.7639 16.000 1.4239 0.08926 0.08374 -0.0759 0.0023 0.7665 16.250 1.4192 0.09336 0.08797 -0.0767 0.0022 0.7692 16.500 1.4144 0.09754 0.09226 -0.0776 0.0022 0.7720 16.750 1.4090 0.10187 0.09671 -0.0787 0.0022 0.7747 17.000 1.4031 0.10629 0.10126 -0.0798 0.0022 0.7774 17.250 1.3959 0.11101 0.10610 -0.0812 0.0021 0.7804 17.500 1.3890 0.11576 0.11098 -0.0827 0.0021 0.7836 17.750 1.3817 0.12063 0.11597 -0.0844 0.0021 0.7871 18.250 1.3673 0.13057 0.12616 -0.0883 0.0020 0.7947 18.500 1.3587 0.13587 0.13159 -0.0906 0.0020 0.7992 18.750 1.3516 0.14100 0.13684 -0.0930 0.0019 0.8041