XFOIL Version 6.94 Calculated polar for: manu01 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 -0.1732 0.05404 0.04852 0.0168 0.7961 0.8262 -2.750 -0.2426 0.05276 0.04745 0.0295 0.8321 0.8084 -2.500 -0.2791 0.05033 0.04509 0.0351 0.8238 0.7844 -2.250 -0.2572 0.04796 0.04239 0.0252 0.5859 0.7101 -2.000 -0.3253 0.04903 0.04404 0.0418 0.6316 0.7268 -1.500 0.0289 0.04660 0.03532 -0.0394 0.4120 0.2888 -1.250 0.0712 0.04524 0.03326 -0.0424 0.4040 0.2825 -1.000 0.1201 0.04380 0.03127 -0.0466 0.3969 0.2808 -0.750 0.1729 0.04257 0.02947 -0.0514 0.3916 0.2823 -0.500 0.2332 0.04147 0.02803 -0.0578 0.3868 0.2914 -0.250 0.3050 0.04077 0.02685 -0.0667 0.3822 0.3141 0.000 0.4001 0.03983 0.02545 -0.0804 0.3778 0.3574 0.250 0.5117 0.03785 0.02411 -0.0977 0.3747 0.5402 0.500 0.6496 0.03781 0.02506 -0.1218 0.3726 1.0001 0.750 0.6936 0.03902 0.02590 -0.1253 0.3720 1.0001 1.000 0.7292 0.04008 0.02676 -0.1269 0.3714 1.0001 1.250 0.7607 0.04114 0.02769 -0.1277 0.3706 1.0001 1.500 0.7898 0.04226 0.02872 -0.1280 0.3699 1.0001 1.750 0.8170 0.04347 0.02987 -0.1279 0.3696 1.0001 2.000 0.8432 0.04482 0.03118 -0.1278 0.3700 1.0001 2.250 0.8695 0.04637 0.03266 -0.1277 0.3712 1.0001 2.500 0.8925 0.04758 0.03398 -0.1269 0.3734 1.0001 2.750 0.9113 0.04872 0.03537 -0.1255 0.3773 1.0001 3.000 0.9302 0.05030 0.03715 -0.1243 0.3815 1.0001 3.250 0.9475 0.05212 0.03910 -0.1230 0.3856 1.0001 3.500 0.9653 0.05413 0.04119 -0.1218 0.3895 1.0001 3.750 0.9854 0.05639 0.04346 -0.1210 0.3930 1.0001 4.000 0.9935 0.05802 0.04544 -0.1185 0.4000 1.0001 4.250 0.9961 0.06055 0.04826 -0.1157 0.4085 1.0001 4.500 1.0114 0.06332 0.05110 -0.1148 0.4146 1.0001 4.750 0.9942 0.06640 0.05458 -0.1102 0.4277 1.0001 5.000 1.0040 0.06992 0.05820 -0.1094 0.4375 1.0001