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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0676 0.07758 0.05592 -0.0198 0.9999 1.0001 -2.750 -0.0694 0.07561 0.05419 -0.0186 0.9999 1.0001 -2.500 -0.0726 0.07361 0.05246 -0.0173 0.9999 1.0001 -2.250 -0.0772 0.07157 0.05070 -0.0157 0.9999 1.0001 -2.000 -0.0832 0.06950 0.04889 -0.0140 0.9999 1.0001 -1.750 -0.0884 0.06746 0.04703 -0.0125 0.9999 1.0001 -1.500 -0.0888 0.06562 0.04517 -0.0120 0.9999 1.0001 -1.250 -0.0798 0.06420 0.04347 -0.0132 0.9999 1.0001 -1.000 -0.0577 0.06336 0.04205 -0.0166 0.9999 1.0001 -0.750 -0.0230 0.06313 0.04092 -0.0219 0.9999 1.0001 -0.500 0.0195 0.06340 0.04015 -0.0280 0.9999 1.0001 -0.250 0.0629 0.06400 0.03965 -0.0334 0.9999 1.0001 0.000 0.1022 0.06469 0.03939 -0.0372 0.9999 1.0001 0.250 0.1369 0.06541 0.03929 -0.0397 0.9999 1.0001 0.500 0.1678 0.06613 0.03945 -0.0413 0.9999 1.0001 0.750 0.1957 0.06692 0.03986 -0.0424 0.9999 1.0001 1.000 0.2210 0.06782 0.04053 -0.0432 0.9999 1.0001 1.250 0.2431 0.06895 0.04160 -0.0440 0.9999 1.0001 1.500 0.2604 0.07050 0.04323 -0.0447 0.9999 1.0001 1.750 0.2690 0.07284 0.04572 -0.0454 0.9999 1.0001 2.000 0.2667 0.07617 0.04908 -0.0461 0.9999 1.0001 2.250 0.2610 0.07989 0.05263 -0.0468 0.9999 1.0001 2.500 0.2589 0.08339 0.05585 -0.0477 0.9999 1.0001 2.750 0.2608 0.08658 0.05872 -0.0485 0.9999 1.0001 3.000 0.2650 0.08960 0.06141 -0.0494 0.9999 1.0001 3.250 0.2709 0.09247 0.06396 -0.0503 0.9999 1.0001 3.750 0.2859 0.09800 0.06884 -0.0519 0.9999 1.0001 4.000 0.2945 0.10068 0.07121 -0.0527 0.9999 1.0001 4.250 0.3036 0.10332 0.07356 -0.0535 0.9999 1.0001 4.500 0.3130 0.10593 0.07589 -0.0542 0.9999 1.0001 4.750 0.3229 0.10852 0.07820 -0.0550 0.9999 1.0001 5.000 0.3329 0.11110 0.08053 -0.0557 0.9999 1.0001