XFOIL Version 6.94 Calculated polar for: manu04 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 -0.4705 0.06155 0.05435 0.0624 0.9999 0.7429 -2.750 -0.4574 0.06056 0.05318 0.0509 0.9999 0.6598 -2.500 -0.3856 0.06001 0.05209 0.0276 0.9999 0.5303 -2.250 -0.3220 0.05875 0.05018 0.0118 0.9999 0.4362 -2.000 -0.2794 0.05669 0.04774 0.0052 0.9999 0.3930 -1.750 -0.2472 0.05432 0.04528 0.0029 0.9999 0.3773 -1.500 -0.2134 0.05247 0.04331 0.0007 0.9999 0.3678 -1.250 -0.1791 0.05047 0.04133 -0.0012 0.9999 0.3605 -1.000 -0.0069 0.04367 0.03392 -0.0277 0.8298 0.3484 -0.750 0.1426 0.03945 0.02684 -0.0470 0.5297 0.3694 -0.500 0.2050 0.03905 0.02543 -0.0539 0.5042 0.3921 -0.250 0.2785 0.03818 0.02402 -0.0629 0.4869 0.4315 0.000 0.3596 0.03595 0.02222 -0.0730 0.4750 0.5844 0.250 0.5020 0.03590 0.02202 -0.0974 0.4596 1.0001 0.500 0.5468 0.03675 0.02240 -0.1009 0.4532 1.0001 0.750 0.5890 0.03770 0.02293 -0.1037 0.4464 1.0001 1.000 0.6313 0.03878 0.02363 -0.1067 0.4411 1.0001 1.250 0.6714 0.03984 0.02449 -0.1093 0.4381 1.0001 1.500 0.7080 0.04081 0.02540 -0.1112 0.4371 1.0001 1.750 0.7431 0.04188 0.02645 -0.1129 0.4367 1.0001 2.000 0.7764 0.04303 0.02763 -0.1143 0.4367 1.0001 2.250 0.8078 0.04429 0.02894 -0.1153 0.4372 1.0001 2.500 0.8374 0.04566 0.03036 -0.1161 0.4380 1.0001 2.750 0.8655 0.04717 0.03192 -0.1166 0.4391 1.0001 3.000 0.8861 0.04848 0.03347 -0.1159 0.4413 1.0001 3.250 0.9043 0.04998 0.03521 -0.1150 0.4443 1.0001 3.500 0.9216 0.05175 0.03724 -0.1141 0.4477 1.0001 3.750 0.9373 0.05375 0.03942 -0.1130 0.4515 1.0001 4.000 0.9531 0.05596 0.04179 -0.1122 0.4555 1.0001 4.250 0.9710 0.05832 0.04423 -0.1116 0.4591 1.0001 4.500 0.9777 0.06080 0.04694 -0.1098 0.4634 1.0001 4.750 0.9617 0.06421 0.05069 -0.1058 0.4710 1.0001 5.000 0.9583 0.06783 0.05443 -0.1035 0.4772 1.0001