XFOIL Version 6.94 Calculated polar for: GOE 406c 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.3576 0.03996 0.03439 -0.0501 0.5614 0.9692 -2.750 0.3004 0.04135 0.03607 -0.0370 0.5764 0.9282 -2.250 0.2036 0.04163 0.03669 -0.0182 0.5657 0.8362 -1.750 0.4318 0.03861 0.02872 -0.1035 0.4196 0.2348 -1.500 0.4684 0.03780 0.02725 -0.1048 0.4066 0.2117 -1.250 0.5031 0.03677 0.02579 -0.1058 0.3972 0.2005 -1.000 0.5419 0.03674 0.02507 -0.1069 0.3890 0.1918 -0.750 0.5784 0.03607 0.02400 -0.1080 0.3823 0.1893 -0.500 0.6144 0.03551 0.02324 -0.1089 0.3760 0.1861 -0.250 0.6502 0.03525 0.02265 -0.1095 0.3696 0.1839 0.000 0.6874 0.03534 0.02224 -0.1102 0.3635 0.1840 0.250 0.7195 0.03538 0.02211 -0.1101 0.3589 0.1875 0.500 0.7508 0.03544 0.02211 -0.1099 0.3552 0.1952 0.750 0.7816 0.03560 0.02229 -0.1099 0.3522 0.2071 1.000 0.8126 0.03590 0.02260 -0.1100 0.3496 0.2224 1.250 0.8446 0.03623 0.02310 -0.1105 0.3472 0.2626 1.500 0.8658 0.03506 0.02343 -0.1079 0.3455 1.0001 1.750 0.8965 0.03650 0.02421 -0.1077 0.3434 1.0001 2.000 0.9240 0.03798 0.02546 -0.1076 0.3414 1.0001 2.250 0.9487 0.03929 0.02677 -0.1072 0.3394 1.0001 2.500 0.9725 0.04074 0.02825 -0.1068 0.3373 1.0001 2.750 0.9955 0.04234 0.02990 -0.1064 0.3355 1.0001 3.000 1.0176 0.04412 0.03174 -0.1059 0.3348 1.0001 3.250 1.0382 0.04608 0.03379 -0.1054 0.3347 1.0001 3.500 1.0569 0.04823 0.03607 -0.1047 0.3351 1.0001 3.750 1.0736 0.05062 0.03861 -0.1039 0.3359 1.0001 4.000 1.0886 0.05325 0.04138 -0.1031 0.3370 1.0001 4.250 1.1011 0.05617 0.04444 -0.1021 0.3385 1.0001 4.500 1.1115 0.05935 0.04777 -0.1011 0.3399 1.0001 4.750 1.1213 0.06277 0.05129 -0.1002 0.3411 1.0001 5.000 1.1293 0.06642 0.05508 -0.0993 0.3425 1.0001