XFOIL Version 6.94 Calculated polar for: manu4/33 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 -0.2287 0.05798 0.05055 0.0130 0.4111 0.6784 -2.750 -0.2168 0.05493 0.04735 0.0163 0.4027 0.7034 -2.500 -0.2124 0.05240 0.04478 0.0191 0.3978 0.7197 -2.250 -0.2050 0.05009 0.04241 0.0204 0.3930 0.7284 -2.000 -0.1852 0.04802 0.04017 0.0184 0.3872 0.7236 -1.750 -0.1026 0.04956 0.04083 -0.0034 0.3770 0.6225 -1.500 0.0708 0.04820 0.03684 -0.0466 0.3646 0.2593 -1.250 0.1249 0.04665 0.03452 -0.0520 0.3579 0.2557 -1.000 0.1867 0.04501 0.03230 -0.0588 0.3541 0.2554 -0.750 0.2416 0.04361 0.03051 -0.0641 0.3518 0.2558 -0.500 0.2989 0.04240 0.02917 -0.0700 0.3496 0.2616 -0.250 0.3705 0.04174 0.02806 -0.0789 0.3478 0.2766 0.000 0.4439 0.04132 0.02746 -0.0884 0.3465 0.3050 0.250 0.5455 0.04026 0.02635 -0.1039 0.3458 0.3975 0.500 0.6083 0.03994 0.02652 -0.1114 0.3461 0.5095 0.750 0.6504 0.04009 0.02723 -0.1144 0.3470 0.5920 1.250 0.7929 0.04258 0.03045 -0.1341 0.3498 1.0001 1.500 0.8231 0.04399 0.03170 -0.1348 0.3507 1.0001 1.750 0.8503 0.04542 0.03301 -0.1348 0.3513 1.0001 2.000 0.8751 0.04688 0.03440 -0.1344 0.3516 1.0001 2.250 0.8984 0.04846 0.03593 -0.1337 0.3525 1.0001 2.500 0.9209 0.05020 0.03761 -0.1329 0.3535 1.0001 2.750 0.9417 0.05156 0.03905 -0.1318 0.3564 1.0001 3.000 0.9543 0.05277 0.04064 -0.1293 0.3637 1.0001 3.250 0.9689 0.05477 0.04278 -0.1275 0.3694 1.0001 3.500 0.9864 0.05706 0.04510 -0.1263 0.3742 1.0001 3.750 1.0095 0.05988 0.04783 -0.1262 0.3781 1.0001 4.000 1.0008 0.06136 0.04995 -0.1213 0.3924 1.0001 4.250 1.0185 0.06436 0.05294 -0.1206 0.3998 1.0001 4.500 0.9992 0.06742 0.05650 -0.1157 0.4186 1.0001 4.750 0.9691 0.07213 0.06159 -0.1108 0.4425 1.0001 5.000 0.9598 0.07804 0.06766 -0.1104 0.4732 1.0001