XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2123 0.05408 0.04663 0.0037 0.3970 0.6365 -2.750 -0.2158 0.05181 0.04430 0.0089 0.3904 0.6661 -2.500 -0.2094 0.04908 0.04153 0.0126 0.3839 0.6892 -2.250 -0.1972 0.04710 0.03942 0.0132 0.3780 0.6957 -2.000 -0.1691 0.04604 0.03801 0.0087 0.3722 0.6809 -1.750 0.0248 0.04553 0.03462 -0.0437 0.3600 0.2487 -1.500 0.0660 0.04347 0.03190 -0.0464 0.3540 0.2371 -1.250 0.1192 0.04167 0.02942 -0.0515 0.3470 0.2318 -1.000 0.1661 0.04030 0.02767 -0.0551 0.3420 0.2328 -0.750 0.2207 0.03931 0.02620 -0.0602 0.3378 0.2405 -0.500 0.2811 0.03838 0.02504 -0.0668 0.3349 0.2489 -0.250 0.3629 0.03771 0.02390 -0.0778 0.3321 0.2645 0.000 0.4515 0.03729 0.02313 -0.0906 0.3298 0.3031 0.250 0.5392 0.03651 0.02270 -0.1034 0.3286 0.4323 0.500 0.5876 0.03668 0.02320 -0.1078 0.3284 0.5019 0.750 0.6259 0.03693 0.02390 -0.1101 0.3289 0.5651 1.000 0.7327 0.03779 0.02575 -0.1279 0.3302 1.0001 1.250 0.7643 0.03887 0.02676 -0.1286 0.3321 1.0001 1.500 0.7935 0.04002 0.02784 -0.1289 0.3339 1.0001 1.750 0.8203 0.04124 0.02903 -0.1287 0.3356 1.0001 2.000 0.8454 0.04253 0.03031 -0.1282 0.3368 1.0001 2.250 0.8687 0.04387 0.03164 -0.1274 0.3375 1.0001 2.500 0.8910 0.04531 0.03310 -0.1265 0.3388 1.0001 2.750 0.9122 0.04683 0.03467 -0.1254 0.3405 1.0001 3.000 0.9326 0.04853 0.03643 -0.1243 0.3429 1.0001 3.250 0.9538 0.05046 0.03836 -0.1235 0.3459 1.0001 3.500 0.9767 0.05285 0.04067 -0.1231 0.3485 1.0001 3.750 0.9835 0.05371 0.04218 -0.1197 0.3595 1.0001 4.000 0.9988 0.05609 0.04467 -0.1183 0.3660 1.0001 4.250 1.0208 0.05898 0.04750 -0.1180 0.3708 1.0001 4.500 1.0161 0.06101 0.05017 -0.1143 0.3886 1.0001 5.000 0.9990 0.06866 0.05858 -0.1091 0.4364 1.0001