XFOIL Version 6.94 Calculated polar for: cr001sm 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 0.5150 0.00725 0.00291 -0.1186 0.8799 1.0000 0.500 0.5756 0.00719 0.00262 -0.1190 0.8574 1.0000 1.000 0.6304 0.00723 0.00247 -0.1181 0.8270 1.0000 1.500 0.6842 0.00735 0.00239 -0.1170 0.7931 1.0000 2.000 0.7359 0.00755 0.00244 -0.1156 0.7563 1.0000 2.500 0.7867 0.00782 0.00254 -0.1141 0.7167 1.0000 3.000 0.8365 0.00817 0.00272 -0.1124 0.6753 1.0000 3.500 0.8852 0.00859 0.00299 -0.1106 0.6314 1.0000 4.000 0.9328 0.00906 0.00331 -0.1087 0.5838 1.0000 5.000 1.0239 0.01028 0.00423 -0.1042 0.4724 1.0000 5.500 1.0664 0.01112 0.00483 -0.1016 0.4017 1.0000 6.000 1.1065 0.01220 0.00562 -0.0988 0.3185 1.0000 6.500 1.1454 0.01349 0.00657 -0.0959 0.2427 1.0000 7.000 1.1847 0.01481 0.00761 -0.0933 0.1778 1.0000 7.500 1.2236 0.01616 0.00881 -0.0906 0.1213 1.0000 8.000 1.2550 0.01823 0.01052 -0.0868 0.0581 1.0000 8.500 1.2802 0.02078 0.01305 -0.0817 0.0330 1.0000 9.000 1.3012 0.02336 0.01579 -0.0761 0.0254 1.0000 9.500 1.3211 0.02578 0.01849 -0.0703 0.0217 1.0000 10.000 1.3391 0.02848 0.02141 -0.0648 0.0189 1.0000 10.500 1.3499 0.03360 0.02682 -0.0593 0.0165 1.0000 11.000 1.3653 0.03710 0.03074 -0.0546 0.0150 1.0000 11.500 1.3739 0.04153 0.03560 -0.0499 0.0139 1.0000 12.000 1.3716 0.04712 0.04167 -0.0453 0.0133 1.0000 12.500 1.3575 0.05401 0.04907 -0.0416 0.0130 1.0000 13.000 1.3335 0.06232 0.05786 -0.0400 0.0128 1.0000 13.500 1.2993 0.07305 0.06911 -0.0415 0.0129 1.0000 14.000 1.2596 0.08627 0.08274 -0.0467 0.0130 1.0000 14.500 1.2147 0.10308 0.09995 -0.0563 0.0133 1.0000 15.000 1.1670 0.12393 0.12115 -0.0699 0.0139 1.0000 15.500 1.1103 0.15206 0.14950 -0.0879 0.0148 1.0000 16.000 1.0704 0.17781 0.17525 -0.1016 0.0163 1.0000