XFOIL Version 6.94 Calculated polar for: AH 79-100 A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4892 0.00735 0.00212 -0.1170 0.7914 0.5709 0.500 0.5438 0.00733 0.00210 -0.1163 0.7697 0.6007 1.000 0.5979 0.00733 0.00212 -0.1156 0.7449 0.6318 1.500 0.6516 0.00734 0.00218 -0.1148 0.7184 0.6683 2.000 0.7038 0.00734 0.00228 -0.1137 0.6872 0.7195 3.000 0.8094 0.00728 0.00252 -0.1117 0.6143 1.0000 3.500 0.8617 0.00768 0.00277 -0.1109 0.5768 1.0000 4.000 0.9138 0.00811 0.00308 -0.1100 0.5402 1.0000 4.500 0.9648 0.00860 0.00344 -0.1090 0.4982 1.0000 5.000 1.0143 0.00918 0.00387 -0.1077 0.4466 1.0000 5.500 1.0626 0.00987 0.00438 -0.1063 0.3928 1.0000 6.000 1.1055 0.01104 0.00512 -0.1042 0.3003 1.0000 6.500 1.1462 0.01248 0.00606 -0.1019 0.2039 1.0000 7.000 1.1862 0.01398 0.00718 -0.0996 0.1293 1.0000 7.500 1.2248 0.01556 0.00842 -0.0971 0.0715 1.0000 8.000 1.2567 0.01771 0.01028 -0.0934 0.0270 1.0000 8.500 1.2900 0.01953 0.01220 -0.0897 0.0207 1.0000 9.000 1.3224 0.02114 0.01396 -0.0861 0.0182 1.0000 9.500 1.3392 0.02339 0.01634 -0.0802 0.0164 1.0000 10.000 1.3523 0.02604 0.01918 -0.0743 0.0157 1.0000 10.500 1.3695 0.02863 0.02197 -0.0697 0.0152 1.0000 11.000 1.3850 0.03170 0.02527 -0.0654 0.0146 1.0000 11.500 1.3999 0.03520 0.02901 -0.0615 0.0141 1.0000 12.000 1.4137 0.03911 0.03317 -0.0582 0.0137 1.0000 12.500 1.4243 0.04339 0.03770 -0.0553 0.0133 1.0000 13.000 1.4305 0.04825 0.04280 -0.0529 0.0128 1.0000 13.500 1.4302 0.05442 0.04930 -0.0510 0.0126 1.0000 14.000 1.4201 0.06209 0.05735 -0.0500 0.0124 1.0000 14.500 1.4010 0.07117 0.06684 -0.0506 0.0124 1.0000 15.000 1.3753 0.08185 0.07793 -0.0535 0.0124 1.0000 15.500 1.3448 0.09451 0.09100 -0.0589 0.0125 1.0000 16.000 1.3072 0.11020 0.10714 -0.0677 0.0127 1.0000 16.500 1.2172 0.14231 0.14000 -0.0896 0.0137 1.0000 17.000 1.1188 0.18778 0.18575 -0.1188 0.0165 1.0000 17.500 0.8692 0.22040 0.21865 -0.1238 0.0258 1.0000 18.000 0.8713 0.22911 0.22737 -0.1279 0.0233 1.0000 18.500 0.8808 0.23499 0.23329 -0.1302 0.0220 1.0000 19.000 0.8875 0.24301 0.24133 -0.1324 0.0210 1.0000 19.500 0.8828 0.25487 0.25318 -0.1390 0.0192 1.0000 20.000 0.8906 0.26158 0.25992 -0.1417 0.0169 1.0000 20.500 0.9019 0.26814 0.26652 -0.1428 0.0160 1.0000 21.000 0.8983 0.28052 0.27888 -0.1491 0.0149 1.0000 21.500 0.9040 0.28909 0.28748 -0.1524 0.0135 1.0000 22.000 0.9110 0.29653 0.29494 -0.1552 0.0127 1.0000 22.500 0.9216 0.30347 0.30190 -0.1563 0.0121 1.0000 23.000 0.9203 0.31542 0.31385 -0.1620 0.0113 1.0000 23.500 0.9258 0.32457 0.32303 -0.1651 0.0103 1.0000 24.000 0.9319 0.33231 0.33079 -0.1679 0.0094 1.0000 24.500 0.9399 0.33996 0.33846 -0.1695 0.0090 1.0000 25.000 0.9431 0.34963 0.34815 -0.1732 0.0089 1.0000 25.500 0.9461 0.36118 0.35971 -0.1773 0.0082 1.0000