XFOIL Version 6.94 Calculated polar for: GOE 406b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1801 0.05703 0.05268 -0.0248 0.8822 0.9105 -2.750 0.1348 0.05698 0.05271 -0.0170 0.8655 0.8556 -2.500 0.0895 0.05677 0.05259 -0.0100 0.8497 0.8108 -2.250 0.0370 0.05654 0.05246 -0.0024 0.8396 0.7794 -1.750 -0.0504 0.05425 0.05033 0.0096 0.8040 0.7171 -1.500 -0.0329 0.05095 0.04666 -0.0065 0.7687 0.5877 -1.250 0.1175 0.04545 0.03925 -0.0391 0.6771 0.3436 -1.000 0.1735 0.04223 0.03520 -0.0418 0.5960 0.2961 -0.750 0.2247 0.04002 0.03192 -0.0440 0.5385 0.2688 -0.500 0.2631 0.03901 0.02995 -0.0446 0.5047 0.2501 -0.250 0.2984 0.03872 0.02866 -0.0446 0.4829 0.2337 0.000 0.3317 0.03770 0.02718 -0.0450 0.4655 0.2267 0.250 0.3721 0.03723 0.02603 -0.0464 0.4506 0.2190 0.500 0.4214 0.03708 0.02506 -0.0493 0.4371 0.2150 0.750 0.4743 0.03687 0.02446 -0.0534 0.4261 0.2183 1.000 0.5369 0.03682 0.02394 -0.0591 0.4179 0.2243 1.250 0.5853 0.03703 0.02403 -0.0624 0.4120 0.2305 1.500 0.6321 0.03716 0.02417 -0.0656 0.4063 0.2447 1.750 0.6764 0.03737 0.02436 -0.0682 0.4009 0.2753 2.000 0.7084 0.03737 0.02510 -0.0692 0.3967 0.3489 2.250 0.8016 0.03825 0.02627 -0.0807 0.3886 1.0000 2.500 0.8233 0.03970 0.02745 -0.0797 0.3857 1.0000 2.750 0.8428 0.04132 0.02894 -0.0786 0.3842 1.0000 3.000 0.8599 0.04311 0.03066 -0.0772 0.3832 1.0000 3.250 0.8730 0.04511 0.03267 -0.0755 0.3828 1.0000 3.500 0.8797 0.04744 0.03509 -0.0732 0.3830 1.0000 3.750 0.8756 0.05034 0.03820 -0.0699 0.3843 1.0000 4.000 0.8578 0.05415 0.04228 -0.0657 0.3864 1.0000 4.250 0.8178 0.05958 0.04804 -0.0606 0.3905 1.0000 4.500 0.7735 0.06591 0.05458 -0.0563 0.3957 1.0000 4.750 0.7560 0.07083 0.05949 -0.0543 0.3987 1.0000