XFOIL Version 6.94 Calculated polar for: NACA 4412 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0099 0.02521 0.01621 -0.0900 0.9188 0.1652 -2.750 0.0391 0.02439 0.01532 -0.0938 0.9144 0.1877 -2.500 0.0663 0.02389 0.01485 -0.0938 0.9023 0.2080 -2.250 0.1129 0.02306 0.01421 -0.0973 0.8975 0.2405 -2.000 0.1415 0.02261 0.01389 -0.0974 0.8862 0.2720 -1.750 0.1868 0.02180 0.01326 -0.1005 0.8807 0.3245 -1.250 0.2552 0.01999 0.01248 -0.1023 0.8641 0.5099 -1.000 0.3083 0.01833 0.01204 -0.1044 0.8611 1.0000 -0.500 0.3787 0.01800 0.01121 -0.1064 0.8425 1.0000 -0.250 0.4041 0.01807 0.01112 -0.1058 0.8297 1.0000 0.000 0.4448 0.01769 0.01056 -0.1074 0.8235 1.0000 0.250 0.4688 0.01781 0.01056 -0.1065 0.8102 1.0000 0.500 0.5082 0.01744 0.01003 -0.1079 0.8040 1.0000 0.750 0.5308 0.01763 0.01013 -0.1068 0.7902 1.0000 1.000 0.5598 0.01763 0.01003 -0.1066 0.7798 1.0000 1.250 0.5909 0.01754 0.00984 -0.1067 0.7701 1.0000 1.500 0.6164 0.01773 0.00995 -0.1061 0.7583 1.0000 1.750 0.6492 0.01760 0.00971 -0.1064 0.7499 1.0000 2.000 0.6726 0.01790 0.00997 -0.1055 0.7375 1.0000 2.250 0.7063 0.01779 0.00974 -0.1060 0.7300 1.0000 2.500 0.7281 0.01818 0.01012 -0.1049 0.7173 1.0000 2.750 0.7572 0.01830 0.01018 -0.1048 0.7083 1.0000 3.000 0.7825 0.01857 0.01042 -0.1042 0.6975 1.0000 3.250 0.8094 0.01882 0.01063 -0.1039 0.6879 1.0000 3.500 0.8362 0.01905 0.01085 -0.1035 0.6781 1.0000 3.750 0.8617 0.01938 0.01117 -0.1030 0.6683 1.0000 4.000 0.8892 0.01960 0.01136 -0.1027 0.6590 1.0000 4.250 0.9136 0.02000 0.01178 -0.1020 0.6490 1.0000 4.500 0.9415 0.02021 0.01198 -0.1018 0.6399 1.0000 4.750 0.9650 0.02066 0.01248 -0.1010 0.6296 1.0000 5.000 0.9936 0.02085 0.01263 -0.1008 0.6206 1.0000