XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2368 0.08621 0.07222 0.0289 0.9999 0.8506 -2.750 -0.2482 0.08360 0.06985 0.0319 0.9999 0.8421 -2.500 -0.2725 0.08098 0.06746 0.0373 0.9999 0.8333 -2.250 -0.2882 0.07824 0.06496 0.0410 0.9999 0.8263 -2.000 -0.3032 0.07550 0.06238 0.0444 0.9999 0.8197 -1.750 -0.2999 0.07285 0.05991 0.0447 0.9999 0.8172 -1.500 -0.2887 0.07037 0.05757 0.0437 0.9999 0.8172 -1.250 -0.2684 0.06811 0.05550 0.0412 0.9999 0.8205 -1.000 -0.2305 0.06614 0.05380 0.0357 0.9999 0.8300 -0.750 -0.1771 0.06453 0.05266 0.0274 0.9999 0.8516 -0.250 -0.0228 0.06404 0.05357 -0.0036 0.9999 1.0001 0.000 -0.0282 0.06456 0.05438 -0.0039 0.9999 1.0001 0.250 -0.0470 0.06726 0.05726 -0.0036 0.9999 1.0001 0.500 -0.0632 0.07038 0.06027 -0.0040 0.9999 1.0001 0.750 -0.0655 0.07304 0.06265 -0.0062 0.9999 1.0001 1.000 -0.0577 0.07560 0.06483 -0.0098 0.9999 1.0001 1.250 -0.0428 0.07829 0.06706 -0.0143 0.9999 1.0001 1.500 -0.0228 0.08113 0.06938 -0.0195 0.9999 1.0001 1.750 0.0021 0.08418 0.07183 -0.0253 0.9999 1.0001 2.000 0.0303 0.08744 0.07442 -0.0315 0.9999 1.0001 2.250 0.0617 0.09093 0.07715 -0.0381 0.9999 1.0001 2.500 0.0936 0.09459 0.07995 -0.0444 0.9999 1.0001 2.750 0.1234 0.09820 0.08271 -0.0497 0.9999 1.0001 3.000 0.1494 0.10163 0.08530 -0.0539 0.9999 1.0001 3.250 0.1714 0.10486 0.08778 -0.0570 0.9999 1.0001 3.500 0.1907 0.10792 0.09016 -0.0593 0.9999 1.0001 3.750 0.2083 0.11086 0.09248 -0.0612 0.9999 1.0001 4.000 0.2244 0.11373 0.09476 -0.0627 0.9999 1.0001 4.250 0.2397 0.11653 0.09704 -0.0639 0.9999 1.0001 4.500 0.2543 0.11931 0.09934 -0.0650 0.9999 1.0001 4.750 0.2684 0.12204 0.10163 -0.0660 0.9999 1.0001 5.000 0.2821 0.12475 0.10392 -0.0669 0.9999 1.0001