XFOIL Version 6.94 Calculated polar for: manu4/38 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 -0.2896 0.06981 0.05402 0.0281 0.9999 0.8352 -2.750 -0.2883 0.06729 0.05163 0.0291 0.9999 0.8368 -2.500 -0.2769 0.06476 0.04920 0.0283 0.9999 0.8429 -2.250 -0.2564 0.06230 0.04682 0.0257 0.9999 0.8535 -2.000 -0.2144 0.05994 0.04450 0.0190 0.9999 0.8749 -1.750 -0.0656 0.05759 0.04194 -0.0110 0.9999 0.9949 -1.500 -0.0707 0.05578 0.04032 -0.0097 0.9999 1.0001 -1.250 -0.0758 0.05384 0.03850 -0.0081 0.9999 1.0001 -1.000 -0.0705 0.05231 0.03682 -0.0085 0.9999 1.0001 -0.750 -0.0471 0.05145 0.03549 -0.0122 0.9999 1.0001 -0.500 -0.0041 0.05134 0.03456 -0.0193 0.9999 1.0001 -0.250 0.0539 0.05194 0.03402 -0.0285 0.9999 1.0001 0.000 0.1089 0.05281 0.03373 -0.0358 0.9999 1.0001 0.250 0.1510 0.05357 0.03369 -0.0396 0.9999 1.0001 0.500 0.1851 0.05427 0.03390 -0.0416 0.9999 1.0001 0.750 0.2140 0.05508 0.03453 -0.0430 0.9999 1.0001 1.000 0.2367 0.05628 0.03579 -0.0441 0.9999 1.0001 1.250 0.2457 0.05858 0.03832 -0.0450 0.9999 1.0001 1.500 0.2329 0.06273 0.04258 -0.0456 0.9999 1.0001 1.750 0.2231 0.06679 0.04645 -0.0464 0.9999 1.0001 2.000 0.2220 0.07019 0.04957 -0.0474 0.9999 1.0001 2.250 0.2253 0.07326 0.05233 -0.0484 0.9999 1.0001 2.500 0.2312 0.07612 0.05487 -0.0494 0.9999 1.0001 2.750 0.2385 0.07887 0.05731 -0.0503 0.9999 1.0001 3.000 0.2469 0.08155 0.05967 -0.0512 0.9999 1.0001 3.250 0.2562 0.08415 0.06197 -0.0520 0.9999 1.0001 3.500 0.2660 0.08674 0.06428 -0.0527 0.9999 1.0001 3.750 0.2761 0.08931 0.06656 -0.0535 0.9999 1.0001 4.000 0.2866 0.09185 0.06884 -0.0542 0.9999 1.0001 4.250 0.2972 0.09441 0.07115 -0.0549 0.9999 1.0001 4.500 0.3082 0.09695 0.07345 -0.0556 0.9999 1.0001 4.750 0.3192 0.09949 0.07576 -0.0563 0.9999 1.0001 5.000 0.3303 0.10204 0.07809 -0.0569 0.9999 1.0001