XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1829 0.06626 0.05816 -0.0003 0.3744 0.5668 -2.750 -0.1766 0.06307 0.05502 0.0044 0.3703 0.6105 -2.500 -0.1708 0.06010 0.05205 0.0081 0.3660 0.6467 -2.250 -0.1540 0.05675 0.04864 0.0097 0.3613 0.6748 -2.000 -0.1345 0.05382 0.04560 0.0097 0.3574 0.6952 -1.750 -0.1036 0.05095 0.04256 0.0072 0.3540 0.7097 -1.500 -0.0370 0.04945 0.04065 -0.0054 0.3505 0.6897 -1.250 0.2167 0.05117 0.03940 -0.0687 0.3451 0.2591 -1.000 0.2710 0.04956 0.03749 -0.0743 0.3439 0.2582 -0.750 0.3247 0.04835 0.03592 -0.0796 0.3429 0.2594 -0.500 0.3757 0.04738 0.03460 -0.0843 0.3425 0.2603 -0.250 0.4243 0.04669 0.03362 -0.0885 0.3425 0.2643 0.000 0.4807 0.04643 0.03319 -0.0944 0.3429 0.2745 0.250 0.5407 0.04650 0.03302 -0.1012 0.3436 0.2965 0.500 0.6079 0.04663 0.03295 -0.1097 0.3449 0.3427 0.750 0.6864 0.04650 0.03314 -0.1207 0.3462 0.4591 1.000 0.7283 0.04719 0.03419 -0.1240 0.3468 0.5248 1.250 0.7611 0.04798 0.03537 -0.1253 0.3470 0.5836 1.500 0.8002 0.04860 0.03661 -0.1283 0.3477 0.6786 1.750 0.8896 0.05151 0.03990 -0.1427 0.3484 1.0001 2.000 0.9107 0.05190 0.04055 -0.1414 0.3532 1.0001 2.250 0.9294 0.05353 0.04236 -0.1401 0.3596 1.0001 2.500 0.9484 0.05559 0.04444 -0.1390 0.3653 1.0001 2.750 0.9688 0.05794 0.04672 -0.1383 0.3701 1.0001 3.000 0.9943 0.06096 0.04952 -0.1386 0.3740 1.0001 3.250 0.9865 0.06184 0.05104 -0.1332 0.3887 1.0001 3.500 1.0028 0.06482 0.05400 -0.1323 0.3966 1.0001 3.750 0.9869 0.06732 0.05697 -0.1269 0.4147 1.0001 4.000 1.0183 0.07131 0.06074 -0.1290 0.4255 1.0001 4.250 1.0067 0.07561 0.06538 -0.1263 0.4521 1.0001 4.750 0.3436 0.11735 0.10810 -0.0842 0.8412 0.4696 5.000 0.3849 0.12094 0.11215 -0.0898 0.8263 0.5422