XFOIL Version 6.94 Calculated polar for: manu4/25 1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.3581 0.06812 0.05846 -0.0714 0.3801 0.9117 -2.750 0.3325 0.06708 0.05755 -0.0651 0.3806 0.8911 -2.500 0.3004 0.06575 0.05635 -0.0579 0.3813 0.8690 -2.250 0.2640 0.06398 0.05475 -0.0502 0.3822 0.8451 -2.000 0.1933 0.06217 0.05319 -0.0356 0.3835 0.8262 -1.750 0.1255 0.06000 0.05125 -0.0217 0.3850 0.8088 -1.500 0.0725 0.05797 0.04932 -0.0117 0.3862 0.7872 -1.250 0.0928 0.05802 0.04917 -0.0187 0.3862 0.7282 -1.000 0.2736 0.06417 0.05410 -0.0648 0.3848 0.4768 -0.750 0.3499 0.06383 0.05322 -0.0783 0.3852 0.3962 -0.500 0.4105 0.06377 0.05260 -0.0869 0.3861 0.3526 -0.250 0.4511 0.06257 0.05134 -0.0902 0.3876 0.3405 0.000 0.4943 0.06246 0.05090 -0.0940 0.3895 0.3284 0.250 0.5347 0.06223 0.05052 -0.0970 0.3916 0.3262 0.500 0.5807 0.06253 0.05053 -0.1013 0.3940 0.3282 0.750 0.6252 0.06297 0.05083 -0.1054 0.3962 0.3385 1.000 0.6706 0.06400 0.05154 -0.1096 0.3979 0.3532 1.250 0.7214 0.06492 0.05236 -0.1153 0.3990 0.3759 1.500 0.7863 0.06601 0.05331 -0.1240 0.4001 0.4326 1.750 0.8407 0.06646 0.05428 -0.1304 0.4022 0.5380 2.000 0.8355 0.06684 0.05530 -0.1253 0.4086 0.5692 2.250 0.8364 0.06880 0.05768 -0.1221 0.4157 0.6080 2.500 0.8582 0.07088 0.06047 -0.1238 0.4236 0.6892 3.000 0.7834 0.07860 0.06863 -0.1072 0.4440 0.6608 3.250 0.7583 0.08383 0.07407 -0.1028 0.4572 0.6676 3.500 0.5681 0.09846 0.08800 -0.0816 0.4953 0.4302 4.250 0.0621 0.12695 0.11827 -0.0392 0.8659 0.3102 4.500 0.0810 0.12864 0.11966 -0.0413 0.8683 0.3091 4.750 0.1176 0.13193 0.12267 -0.0470 0.8649 0.3110 5.000 0.1689 0.13825 0.12866 -0.0556 0.8610 0.3166