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 -0.2066 0.05479 0.04723 -0.0005 0.3954 0.6058 -2.750 -0.2055 0.05256 0.04495 0.0042 0.3865 0.6354 -2.500 -0.2034 0.05064 0.04292 0.0077 0.3795 0.6585 -2.250 -0.1932 0.04822 0.04036 0.0100 0.3738 0.6758 -2.000 -0.1694 0.04632 0.03827 0.0082 0.3678 0.6746 -1.750 0.0228 0.04515 0.03406 -0.0436 0.3558 0.2367 -1.500 0.0673 0.04318 0.03123 -0.0469 0.3512 0.2289 -1.250 0.1148 0.04180 0.02963 -0.0508 0.3471 0.2345 -1.000 0.1659 0.04037 0.02778 -0.0552 0.3437 0.2357 -0.750 0.2233 0.03913 0.02610 -0.0610 0.3405 0.2389 -0.500 0.2878 0.03809 0.02457 -0.0681 0.3379 0.2470 -0.250 0.3670 0.03760 0.02381 -0.0788 0.3354 0.2687 0.000 0.4511 0.03717 0.02314 -0.0906 0.3337 0.3166 0.250 0.5421 0.03623 0.02255 -0.1041 0.3314 0.4539 0.500 0.5893 0.03648 0.02313 -0.1084 0.3295 0.5290 0.750 0.6282 0.03691 0.02399 -0.1109 0.3278 0.6145 1.000 0.7332 0.03811 0.02580 -0.1283 0.3262 1.0001 1.250 0.7650 0.03935 0.02685 -0.1292 0.3268 1.0001 1.500 0.7949 0.04069 0.02802 -0.1297 0.3276 1.0001 1.750 0.8219 0.04163 0.02896 -0.1294 0.3291 1.0001 2.000 0.8456 0.04252 0.02999 -0.1284 0.3320 1.0001 2.250 0.8686 0.04378 0.03138 -0.1275 0.3354 1.0001 2.500 0.8912 0.04525 0.03293 -0.1266 0.3389 1.0001 2.750 0.9128 0.04686 0.03461 -0.1257 0.3423 1.0001 3.000 0.9343 0.04865 0.03643 -0.1248 0.3456 1.0001 3.250 0.9566 0.05075 0.03849 -0.1242 0.3484 1.0001 3.500 0.9705 0.05180 0.03998 -0.1219 0.3565 1.0001 3.750 0.9841 0.05397 0.04238 -0.1201 0.3643 1.0001 4.000 1.0039 0.05650 0.04490 -0.1194 0.3696 1.0001 4.250 1.0064 0.05843 0.04735 -0.1164 0.3846 1.0001 4.500 1.0258 0.06142 0.05037 -0.1160 0.3926 1.0001 4.750 1.0200 0.06442 0.05384 -0.1130 0.4133 1.0001 5.000 1.0164 0.06854 0.05826 -0.1115 0.4380 1.0001