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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0616 0.07949 0.05806 -0.0195 0.9999 1.0001 -2.750 -0.0631 0.07751 0.05633 -0.0184 0.9999 1.0001 -2.500 -0.0657 0.07552 0.05459 -0.0171 0.9999 1.0001 -2.250 -0.0699 0.07347 0.05282 -0.0156 0.9999 1.0001 -2.000 -0.0754 0.07139 0.05102 -0.0139 0.9999 1.0001 -1.750 -0.0812 0.06931 0.04917 -0.0123 0.9999 1.0001 -1.500 -0.0838 0.06736 0.04731 -0.0113 0.9999 1.0001 -1.250 -0.0790 0.06575 0.04554 -0.0117 0.9999 1.0001 -1.000 -0.0623 0.06468 0.04404 -0.0142 0.9999 1.0001 -0.750 -0.0327 0.06422 0.04288 -0.0187 0.9999 1.0001 -0.500 0.0071 0.06432 0.04204 -0.0246 0.9999 1.0001 -0.250 0.0511 0.06485 0.04156 -0.0306 0.9999 1.0001 0.000 0.0932 0.06559 0.04133 -0.0354 0.9999 1.0001 0.250 0.1306 0.06638 0.04128 -0.0387 0.9999 1.0001 0.500 0.1634 0.06721 0.04150 -0.0410 0.9999 1.0001 0.750 0.1922 0.06811 0.04201 -0.0425 0.9999 1.0001 1.000 0.2175 0.06917 0.04285 -0.0436 0.9999 1.0001 1.250 0.2381 0.07055 0.04419 -0.0446 0.9999 1.0001 1.500 0.2514 0.07254 0.04624 -0.0455 0.9999 1.0001 1.750 0.2536 0.07552 0.04929 -0.0463 0.9999 1.0001 2.000 0.2478 0.07925 0.05290 -0.0470 0.9999 1.0001 2.250 0.2442 0.08288 0.05626 -0.0478 0.9999 1.0001 2.500 0.2449 0.08619 0.05924 -0.0487 0.9999 1.0001 2.750 0.2485 0.08929 0.06198 -0.0496 0.9999 1.0001 3.000 0.2541 0.09223 0.06456 -0.0505 0.9999 1.0001 3.250 0.2610 0.09507 0.06705 -0.0514 0.9999 1.0001 3.500 0.2688 0.09783 0.06946 -0.0523 0.9999 1.0001 3.750 0.2774 0.10054 0.07183 -0.0531 0.9999 1.0001 4.000 0.2865 0.10321 0.07416 -0.0539 0.9999 1.0001 4.250 0.2960 0.10583 0.07647 -0.0547 0.9999 1.0001 4.500 0.3060 0.10843 0.07875 -0.0554 0.9999 1.0001 4.750 0.3161 0.11103 0.08106 -0.0562 0.9999 1.0001 5.000 0.3265 0.11360 0.08334 -0.0569 0.9999 1.0001