XFOIL Version 6.94 Calculated polar for: agora05 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 -0.1087 0.03973 0.02890 -0.0242 0.9992 0.4238 -2.750 -0.0861 0.03725 0.02743 -0.0212 0.9992 0.5590 -2.500 -0.0805 0.03456 0.02660 -0.0120 0.9992 0.8040 -2.250 -0.0320 0.03407 0.02507 -0.0192 0.9992 1.0008 -2.000 -0.0048 0.03461 0.02509 -0.0199 0.9992 1.0008 -1.750 0.0204 0.03519 0.02533 -0.0200 0.9992 1.0008 -1.500 0.0449 0.03579 0.02570 -0.0201 0.9992 1.0008 -1.250 0.0691 0.03643 0.02620 -0.0201 0.9992 1.0008 -1.000 0.0932 0.03712 0.02683 -0.0202 0.9992 1.0008 -0.500 0.2819 0.04069 0.03037 -0.0253 0.7212 1.0008 -0.250 0.2823 0.04486 0.03442 -0.0165 0.6339 1.0008 0.000 0.3025 0.04737 0.03687 -0.0144 0.5541 1.0008 0.250 0.3559 0.04602 0.03563 -0.0226 0.4781 1.0008 0.500 0.3773 0.04475 0.03409 -0.0170 0.4591 1.0008 0.750 0.3966 0.04331 0.03229 -0.0107 0.4426 1.0008 1.000 0.4170 0.04238 0.03094 -0.0055 0.4252 1.0008 1.250 0.4373 0.04202 0.03008 -0.0009 0.4098 1.0008 1.500 0.4652 0.04312 0.03097 0.0001 0.3965 1.0008 1.750 0.4921 0.04437 0.03197 0.0014 0.3846 1.0008 2.000 0.5222 0.04633 0.03385 0.0009 0.3720 1.0008 2.250 0.5519 0.04856 0.03603 0.0002 0.3620 1.0008 2.500 0.5854 0.05180 0.03945 -0.0029 0.3565 1.0008 2.750 0.6175 0.05542 0.04322 -0.0062 0.3510 1.0008 3.000 0.6504 0.06089 0.04902 -0.0129 0.3457 1.0008 3.250 0.6787 0.06761 0.05601 -0.0208 0.3443 1.0008 3.500 0.6778 0.08629 0.07527 -0.0464 0.3660 1.0008 4.250 0.5894 0.11554 0.10460 -0.0707 0.4718 1.0008