XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0750 0.04483 0.03496 -0.0525 0.3494 0.2066 -2.750 -0.0592 0.04312 0.03313 -0.0503 0.3399 0.2046 -2.500 -0.0411 0.04146 0.03118 -0.0486 0.3326 0.2009 -2.250 -0.0171 0.03885 0.02780 -0.0478 0.3277 0.1911 -2.000 0.0090 0.03732 0.02601 -0.0473 0.3227 0.1892 -1.750 0.0407 0.03582 0.02410 -0.0478 0.3180 0.1878 -1.500 0.0833 0.03442 0.02215 -0.0505 0.3133 0.1879 -1.250 0.1367 0.03329 0.02041 -0.0554 0.3097 0.1926 -1.000 0.1884 0.03264 0.01965 -0.0601 0.3068 0.2004 -0.750 0.2480 0.03179 0.01840 -0.0664 0.3037 0.2084 -0.500 0.3096 0.03134 0.01768 -0.0732 0.3013 0.2228 -0.250 0.3863 0.03114 0.01738 -0.0837 0.2991 0.2569 0.000 0.4567 0.03077 0.01731 -0.0930 0.2977 0.3603 0.250 0.5010 0.03097 0.01772 -0.0966 0.2968 0.4244 0.500 0.5398 0.03136 0.01823 -0.0990 0.2955 0.4724 0.750 0.5774 0.03184 0.01891 -0.1013 0.2937 0.5174 1.000 0.6932 0.03278 0.02083 -0.1214 0.2912 1.0001 1.250 0.7228 0.03372 0.02163 -0.1217 0.2913 1.0001 1.500 0.7513 0.03478 0.02256 -0.1217 0.2918 1.0001 1.750 0.7788 0.03592 0.02360 -0.1216 0.2924 1.0001 2.000 0.8044 0.03660 0.02435 -0.1209 0.2937 1.0001 2.250 0.8284 0.03748 0.02543 -0.1199 0.2960 1.0001 2.500 0.8522 0.03864 0.02675 -0.1190 0.2991 1.0001 2.750 0.8757 0.03999 0.02821 -0.1182 0.3019 1.0001 3.000 0.8984 0.04146 0.02978 -0.1173 0.3048 1.0001 3.250 0.9209 0.04305 0.03142 -0.1165 0.3076 1.0001 3.500 0.9441 0.04492 0.03330 -0.1159 0.3100 1.0001 3.750 0.9608 0.04610 0.03502 -0.1139 0.3191 1.0001 4.000 0.9795 0.04820 0.03728 -0.1127 0.3254 1.0001 4.250 1.0020 0.05068 0.03972 -0.1122 0.3295 1.0001 4.500 1.0133 0.05275 0.04236 -0.1101 0.3468 1.0001 4.750 0.9552 0.06250 0.05420 -0.1127 0.5246 1.0001