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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4487 0.07111 0.06204 0.0465 0.9999 0.6351 -2.750 -0.4220 0.06913 0.05991 0.0389 0.9999 0.5947 -2.500 -0.3920 0.06704 0.05763 0.0324 0.9999 0.5632 -2.250 -0.3572 0.06547 0.05581 0.0254 0.9999 0.5338 -2.000 -0.3224 0.06348 0.05367 0.0203 0.9999 0.5106 -1.750 -0.2854 0.06168 0.05170 0.0153 0.9999 0.4893 -1.500 -0.2462 0.06015 0.04997 0.0103 0.9999 0.4715 -1.250 -0.2092 0.05805 0.04789 0.0072 0.9999 0.4621 -1.000 -0.1685 0.05643 0.04627 0.0033 0.9999 0.4574 -0.750 -0.1280 0.05497 0.04493 -0.0004 0.9999 0.4587 -0.500 0.0393 0.05049 0.04019 -0.0295 0.8609 0.4697 -0.250 0.1062 0.04787 0.03732 -0.0370 0.7353 0.4827 0.000 0.2302 0.04280 0.03138 -0.0516 0.6434 0.5466 0.250 0.4980 0.03903 0.02533 -0.0976 0.5747 1.0001 0.500 0.5520 0.04052 0.02584 -0.1031 0.5585 1.0001 0.750 0.5981 0.04188 0.02654 -0.1070 0.5477 1.0001 1.000 0.6389 0.04314 0.02738 -0.1100 0.5407 1.0001 1.250 0.6789 0.04445 0.02835 -0.1128 0.5364 1.0001 1.500 0.7169 0.04585 0.02948 -0.1152 0.5333 1.0001 1.750 0.7521 0.04732 0.03075 -0.1173 0.5311 1.0001 2.000 0.7803 0.04879 0.03222 -0.1182 0.5302 1.0001 2.250 0.8060 0.05041 0.03387 -0.1187 0.5299 1.0001 2.500 0.8296 0.05218 0.03568 -0.1190 0.5301 1.0001 2.750 0.8512 0.05413 0.03768 -0.1191 0.5309 1.0001 3.000 0.8640 0.05632 0.04005 -0.1181 0.5326 1.0001 3.250 0.8642 0.05908 0.04306 -0.1157 0.5357 1.0001 3.500 0.8564 0.06256 0.04678 -0.1129 0.5399 1.0001 3.750 0.8434 0.06673 0.05111 -0.1100 0.5453 1.0001 4.000 0.8362 0.07102 0.05545 -0.1082 0.5507 1.0001 4.250 0.7587 0.07937 0.06401 -0.1012 0.5639 1.0001 4.500 0.7410 0.08498 0.06957 -0.1002 0.5738 1.0001 4.750 0.6824 0.09338 0.07798 -0.0981 0.5929 1.0001 5.000 0.6523 0.09938 0.08395 -0.0974 0.6107 1.0001