XFOIL Version 6.94 Calculated polar for: FX 62-K-153/20 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 0.5992 0.01684 0.01148 -0.1467 0.7875 0.7324 0.500 0.6715 0.01653 0.01106 -0.1507 0.7828 0.7396 1.000 0.7223 0.01653 0.01109 -0.1498 0.7753 0.7442 1.500 0.7724 0.01655 0.01116 -0.1492 0.7669 0.7496 2.000 0.8405 0.01635 0.01093 -0.1523 0.7618 0.7562 2.500 0.9055 0.01627 0.01089 -0.1548 0.7563 0.7608 3.000 0.9467 0.01631 0.01108 -0.1522 0.7456 0.7657 3.500 1.0225 0.01545 0.01020 -0.1559 0.7375 0.7704 4.000 1.0702 0.01529 0.01017 -0.1548 0.7241 0.7760 4.500 1.1406 0.01471 0.00963 -0.1578 0.7147 0.7809 5.000 1.1853 0.01423 0.00931 -0.1555 0.6973 0.7854 5.500 1.2335 0.01373 0.00887 -0.1538 0.6741 0.7902 6.000 1.2629 0.01350 0.00859 -0.1484 0.6260 0.7965 6.500 1.2867 0.01382 0.00875 -0.1424 0.5740 0.8013 7.000 1.3033 0.01468 0.00945 -0.1353 0.5214 0.8062 7.500 1.2971 0.01654 0.01093 -0.1251 0.4413 0.8123 8.000 1.2787 0.01966 0.01349 -0.1146 0.3410 0.8185 8.500 1.2660 0.02315 0.01651 -0.1063 0.2517 0.8234 9.000 1.2714 0.02622 0.01928 -0.1012 0.1855 0.8285 9.500 1.2817 0.02939 0.02218 -0.0972 0.1286 0.8339 10.000 1.2993 0.03231 0.02498 -0.0944 0.0947 0.8395 10.500 1.3158 0.03525 0.02785 -0.0914 0.0672 0.8448 11.000 1.3347 0.03825 0.03084 -0.0891 0.0480 0.8504 11.500 1.3562 0.04132 0.03399 -0.0874 0.0361 0.8563 12.000 1.3745 0.04466 0.03741 -0.0856 0.0258 0.8627 12.500 1.3912 0.04833 0.04118 -0.0838 0.0173 0.8702 13.000 1.4047 0.05273 0.04567 -0.0825 0.0076 0.8773 13.500 1.4119 0.05772 0.05085 -0.0808 0.0038 0.8853 14.000 1.4212 0.06285 0.05625 -0.0797 0.0030 0.8952 14.500 1.4253 0.06835 0.06206 -0.0785 0.0026 0.9092 15.500 1.4209 0.08166 0.07606 -0.0776 0.0023 1.0000 16.000 1.4189 0.09026 0.08499 -0.0798 0.0023 1.0000 16.500 1.4124 0.09981 0.09488 -0.0828 0.0022 1.0000 17.000 1.4026 0.11023 0.10563 -0.0867 0.0022 1.0000 17.500 1.3901 0.12142 0.11714 -0.0918 0.0022 1.0000 18.000 1.3755 0.13325 0.12929 -0.0978 0.0022 1.0000 18.500 1.3611 0.14535 0.14170 -0.1048 0.0022 1.0000 19.000 1.3461 0.15785 0.15449 -0.1125 0.0022 1.0000 19.500 1.3308 0.17073 0.16765 -0.1211 0.0022 1.0000 20.000 1.3153 0.18403 0.18122 -0.1303 0.0022 1.0000 20.500 1.2999 0.19772 0.19515 -0.1400 0.0023 1.0000 21.000 1.2843 0.21203 0.20968 -0.1501 0.0023 1.0000 21.500 1.2689 0.22690 0.22475 -0.1605 0.0024 1.0000 22.000 1.2502 0.24430 0.24236 -0.1720 0.0025 1.0000 23.500 0.9568 0.35271 0.35137 -0.1860 0.0120 0.8770 24.000 0.9609 0.36317 0.36185 -0.1889 0.0110 0.8812 24.500 0.9651 0.37200 0.37072 -0.1914 0.0104 0.8868 25.000 0.9728 0.38045 0.37920 -0.1936 0.0101 0.8930 25.500 0.9766 0.39463 0.39339 -0.1983 0.0099 0.8968 26.000 0.9773 0.40403 0.40281 -0.2003 0.0093 0.9029 26.500 0.9803 0.41360 0.41241 -0.2028 0.0087 0.9100 27.000 0.9832 0.42273 0.42156 -0.2051 0.0082 0.9180 27.500 0.9825 0.42867 0.42751 -0.2059 0.0079 0.9305