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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3929 0.06573 0.05898 0.0594 0.9999 0.8193 -2.750 -0.4491 0.06307 0.05646 0.0682 0.9999 0.7923 -2.500 -0.4666 0.06092 0.05433 0.0665 0.9999 0.7397 -2.250 -0.4147 0.06117 0.05427 0.0461 0.9999 0.6255 -2.000 -0.3273 0.06126 0.05388 0.0218 0.9999 0.4893 -1.500 -0.0804 0.05064 0.04245 -0.0212 0.8022 0.3698 -1.250 0.0506 0.04562 0.03469 -0.0383 0.5295 0.3512 -1.000 0.0970 0.04490 0.03299 -0.0423 0.5034 0.3504 -0.750 0.1500 0.04436 0.03153 -0.0475 0.4834 0.3593 -0.500 0.2097 0.04353 0.03008 -0.0539 0.4676 0.3729 -0.250 0.2809 0.04276 0.02864 -0.0627 0.4568 0.3922 0.000 0.3612 0.04173 0.02725 -0.0731 0.4485 0.4439 0.250 0.4629 0.03848 0.02500 -0.0877 0.4415 0.6795 0.500 0.5943 0.03975 0.02594 -0.1102 0.4360 1.0001 0.750 0.6422 0.04079 0.02660 -0.1145 0.4345 1.0001 1.000 0.6851 0.04183 0.02739 -0.1177 0.4337 1.0001 1.250 0.7242 0.04292 0.02832 -0.1202 0.4335 1.0001 1.500 0.7604 0.04407 0.02937 -0.1221 0.4339 1.0001 1.750 0.7935 0.04531 0.03055 -0.1235 0.4346 1.0001 2.000 0.8237 0.04664 0.03188 -0.1243 0.4352 1.0001 2.250 0.8511 0.04806 0.03331 -0.1247 0.4357 1.0001 2.500 0.8758 0.04958 0.03485 -0.1246 0.4361 1.0001 2.750 0.8977 0.05120 0.03653 -0.1240 0.4364 1.0001 3.000 0.9184 0.05295 0.03834 -0.1234 0.4372 1.0001 3.250 0.9372 0.05483 0.04033 -0.1225 0.4379 1.0001 3.500 0.9562 0.05690 0.04245 -0.1217 0.4393 1.0001 3.750 0.9647 0.05878 0.04459 -0.1195 0.4431 1.0001 4.000 0.9631 0.06125 0.04737 -0.1163 0.4490 1.0001 4.250 0.9652 0.06413 0.05042 -0.1139 0.4549 1.0001 4.500 0.9763 0.06708 0.05342 -0.1128 0.4601 1.0001 4.750 0.9451 0.07124 0.05789 -0.1072 0.4695 1.0001 5.000 0.9169 0.07639 0.06318 -0.1031 0.4807 1.0001