XFOIL Version 6.94 Calculated polar for: manu4/36 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 -0.2098 0.05238 0.04510 0.0009 0.4189 0.6380 -2.750 -0.2103 0.05010 0.04270 0.0058 0.4057 0.6601 -2.500 -0.2067 0.04832 0.04078 0.0086 0.3972 0.6761 -2.250 -0.1925 0.04591 0.03825 0.0098 0.3883 0.6826 -2.000 -0.1028 0.04784 0.03915 -0.0153 0.3763 0.5376 -1.750 0.0209 0.04329 0.03193 -0.0436 0.3678 0.2307 -1.500 0.0573 0.04180 0.02994 -0.0452 0.3612 0.2304 -1.250 0.1017 0.04043 0.02803 -0.0484 0.3548 0.2329 -1.000 0.1471 0.03914 0.02646 -0.0518 0.3483 0.2356 -0.750 0.2004 0.03806 0.02496 -0.0567 0.3421 0.2397 -0.500 0.2725 0.03716 0.02339 -0.0654 0.3379 0.2501 -0.250 0.3587 0.03668 0.02265 -0.0777 0.3349 0.2771 0.000 0.4568 0.03562 0.02157 -0.0924 0.3328 0.3588 0.250 0.5269 0.03497 0.02141 -0.1014 0.3317 0.4834 0.500 0.5725 0.03499 0.02191 -0.1051 0.3313 0.5606 0.750 0.6863 0.03572 0.02355 -0.1244 0.3314 1.0001 1.000 0.7200 0.03682 0.02443 -0.1256 0.3321 1.0001 1.250 0.7512 0.03799 0.02542 -0.1263 0.3331 1.0001 1.500 0.7804 0.03918 0.02648 -0.1266 0.3340 1.0001 1.750 0.8077 0.04043 0.02765 -0.1265 0.3345 1.0001 2.000 0.8333 0.04169 0.02885 -0.1261 0.3349 1.0001 2.250 0.8580 0.04306 0.03018 -0.1255 0.3354 1.0001 2.500 0.8812 0.04447 0.03160 -0.1248 0.3356 1.0001 2.750 0.9038 0.04580 0.03301 -0.1239 0.3368 1.0001 3.000 0.9219 0.04667 0.03421 -0.1220 0.3407 1.0001 3.250 0.9401 0.04828 0.03607 -0.1206 0.3459 1.0001 3.500 0.9582 0.05015 0.03808 -0.1192 0.3504 1.0001 3.750 0.9779 0.05223 0.04020 -0.1182 0.3542 1.0001 4.000 1.0009 0.05479 0.04272 -0.1179 0.3574 1.0001 4.250 1.0026 0.05615 0.04476 -0.1144 0.3709 1.0001 4.500 1.0212 0.05887 0.04750 -0.1136 0.3771 1.0001 4.750 1.0186 0.06139 0.05054 -0.1106 0.3944 1.0001