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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5286 0.00566 0.00187 -0.1201 0.8468 1.0000 0.500 0.5824 0.00578 0.00175 -0.1192 0.8167 1.0000 1.000 0.6346 0.00595 0.00170 -0.1179 0.7818 1.0000 1.500 0.6859 0.00618 0.00172 -0.1165 0.7452 1.0000 2.000 0.7363 0.00645 0.00182 -0.1151 0.7061 1.0000 2.500 0.7860 0.00679 0.00195 -0.1135 0.6649 1.0000 3.000 0.8353 0.00715 0.00215 -0.1119 0.6222 1.0000 3.500 0.8838 0.00759 0.00243 -0.1102 0.5772 1.0000 4.000 0.9317 0.00807 0.00275 -0.1084 0.5292 1.0000 4.500 0.9785 0.00865 0.00316 -0.1065 0.4735 1.0000 5.000 1.0240 0.00933 0.00363 -0.1045 0.4104 1.0000 5.500 1.0675 0.01023 0.00421 -0.1022 0.3336 1.0000 6.000 1.1095 0.01133 0.00498 -0.0999 0.2561 1.0000 6.500 1.1513 0.01247 0.00583 -0.0975 0.1885 1.0000 7.000 1.1939 0.01354 0.00671 -0.0954 0.1394 1.0000 7.500 1.2335 0.01487 0.00782 -0.0928 0.0866 1.0000 8.000 1.2667 0.01677 0.00942 -0.0891 0.0368 1.0000 8.500 1.2986 0.01871 0.01145 -0.0850 0.0232 1.0000 9.000 1.3304 0.02043 0.01334 -0.0810 0.0181 1.0000 10.000 1.3685 0.02496 0.01835 -0.0692 0.0132 1.0000 10.500 1.3835 0.02739 0.02093 -0.0636 0.0113 1.0000 11.000 1.3790 0.03274 0.02668 -0.0563 0.0102 1.0000 11.500 1.3920 0.03590 0.03017 -0.0518 0.0097 1.0000 12.000 1.3962 0.04032 0.03495 -0.0473 0.0092 1.0000 12.500 1.3913 0.04601 0.04110 -0.0434 0.0088 1.0000 13.000 1.3784 0.05281 0.04832 -0.0409 0.0085 1.0000 13.500 1.3556 0.06155 0.05750 -0.0404 0.0084 1.0000 14.000 1.3259 0.07228 0.06866 -0.0427 0.0083 1.0000 14.500 1.2888 0.08595 0.08274 -0.0484 0.0084 1.0000 15.000 1.2485 0.10247 0.09965 -0.0576 0.0086 1.0000 15.500 1.2066 0.12193 0.11944 -0.0698 0.0088 1.0000 16.000 1.1612 0.14531 0.14306 -0.0847 0.0091 1.0000 16.500 1.0948 0.17949 0.17739 -0.1045 0.0101 1.0000