XFOIL Version 6.96 Calculated polar for: FX 60-177 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.250 -0.5769 0.10180 0.09624 -0.0666 1.0001 0.0157 -14.000 -0.6026 0.09440 0.08874 -0.0695 1.0001 0.0155 -13.750 -0.6279 0.08795 0.08217 -0.0717 1.0001 0.0154 -13.500 -0.6524 0.08234 0.07643 -0.0729 1.0001 0.0153 -13.250 -0.6760 0.07748 0.07141 -0.0733 1.0001 0.0153 -13.000 -0.6984 0.07322 0.06699 -0.0730 1.0001 0.0152 -12.750 -0.7195 0.06944 0.06304 -0.0720 1.0001 0.0152 -12.500 -0.7391 0.06605 0.05947 -0.0704 1.0001 0.0152 -12.250 -0.7568 0.06301 0.05625 -0.0684 1.0001 0.0153 -12.000 -0.7728 0.06024 0.05329 -0.0661 1.0001 0.0153 -11.750 -0.7804 0.05716 0.04995 -0.0651 0.9988 0.0155 -11.500 -0.7782 0.05396 0.04640 -0.0654 0.9960 0.0158 -11.250 -0.7704 0.05156 0.04384 -0.0658 0.9934 0.0161 -11.000 -0.7617 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0.8437 8.500 1.0395 0.02851 0.01974 -0.0449 0.3077 0.8463 8.750 1.0592 0.02924 0.02049 -0.0447 0.2946 0.8490 9.000 1.0742 0.02996 0.02120 -0.0434 0.2830 0.8514 9.250 1.0881 0.03065 0.02196 -0.0420 0.2708 0.8540 9.500 1.1031 0.03141 0.02278 -0.0408 0.2596 0.8566 9.750 1.1182 0.03226 0.02365 -0.0399 0.2490 0.8591 10.000 1.1337 0.03318 0.02461 -0.0392 0.2382 0.8616 10.250 1.1506 0.03413 0.02564 -0.0388 0.2280 0.8640 10.500 1.1630 0.03512 0.02662 -0.0376 0.2197 0.8662 10.750 1.1759 0.03599 0.02766 -0.0363 0.2106 0.8687 11.000 1.1870 0.03704 0.02877 -0.0350 0.2023 0.8715 11.250 1.1987 0.03816 0.02999 -0.0340 0.1932 0.8744 11.500 1.2106 0.03939 0.03129 -0.0332 0.1850 0.8771 11.750 1.2219 0.04077 0.03272 -0.0326 0.1771 0.8796 12.000 1.2306 0.04204 0.03410 -0.0313 0.1698 0.8819 12.250 1.2382 0.04343 0.03559 -0.0301 0.1624 0.8842 12.500 1.2462 0.04491 0.03719 -0.0290 0.1555 0.8868 12.750 1.2538 0.04650 0.03892 -0.0282 0.1485 0.8896 13.000 1.2610 0.04827 0.04075 -0.0275 0.1427 0.8924 13.250 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