XFOIL Version 6.96 Calculated polar for: WORTMANN FX 79-K-144/17 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.750 -1.0203 0.05612 0.05275 -0.0621 0.7714 0.0046 -15.500 -1.0465 0.04832 0.04474 -0.0674 0.7711 0.0045 -15.250 -1.0621 0.04345 0.03970 -0.0694 0.7707 0.0046 -15.000 -1.0718 0.03988 0.03597 -0.0699 0.7703 0.0046 -14.750 -1.0772 0.03705 0.03300 -0.0695 0.7699 0.0046 -14.500 -1.0793 0.03468 0.03048 -0.0687 0.7695 0.0046 -14.250 -1.0785 0.03265 0.02831 -0.0676 0.7691 0.0047 -14.000 -1.0749 0.03088 0.02641 -0.0664 0.7687 0.0047 -13.750 -1.0690 0.02930 0.02470 -0.0652 0.7683 0.0048 -13.500 -1.0607 0.02789 0.02317 -0.0640 0.7680 0.0049 -13.250 -1.0509 0.02661 0.02176 -0.0628 0.7676 0.0049 -13.000 -1.0415 0.02553 0.02057 -0.0611 0.7672 0.0050 -12.750 -1.0275 0.02449 0.01941 -0.0600 0.7667 0.0050 -12.500 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17.000 1.4218 0.07537 0.07080 -0.0333 0.0039 0.8860 17.500 1.4146 0.08286 0.07852 -0.0341 0.0038 0.8891 17.750 1.4099 0.08686 0.08264 -0.0348 0.0038 0.8910 18.000 1.4039 0.09113 0.08703 -0.0356 0.0038 0.8930 18.250 1.3975 0.09557 0.09159 -0.0366 0.0037 0.8952 18.500 1.3900 0.10024 0.09638 -0.0379 0.0037 0.8974 18.750 1.3822 0.10504 0.10130 -0.0393 0.0037 0.8998