XFOIL Version 6.94 Calculated polar for: nacb632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1123 0.03152 0.02235 -0.0017 1.0000 0.9406 -2.750 -0.0802 0.03054 0.02119 -0.0046 1.0000 0.9556 -2.500 -0.0473 0.02962 0.02013 -0.0080 1.0000 0.9698 -2.250 -0.0122 0.02875 0.01913 -0.0119 1.0000 0.9834 -2.000 0.0287 0.02787 0.01813 -0.0169 1.0000 0.9973 -1.750 0.0329 0.02769 0.01792 -0.0157 1.0000 1.0000 -1.500 0.0265 0.02766 0.01788 -0.0126 1.0000 1.0000 -1.250 0.0198 0.02759 0.01781 -0.0096 1.0000 1.0000 -1.000 0.0128 0.02748 0.01769 -0.0065 1.0000 1.0000 -0.750 0.0055 0.02732 0.01753 -0.0035 1.0000 1.0000 -0.500 -0.0024 0.02711 0.01732 -0.0004 1.0000 1.0000 -0.250 -0.0110 0.02685 0.01705 0.0028 1.0000 1.0000 0.000 -0.0205 0.02652 0.01672 0.0061 1.0000 1.0000 0.250 -0.0307 0.02613 0.01632 0.0096 1.0000 1.0000 0.500 -0.0405 0.02569 0.01588 0.0129 1.0000 1.0000 0.750 -0.0461 0.02534 0.01549 0.0157 1.0000 1.0000 1.000 -0.0401 0.02530 0.01539 0.0167 1.0000 1.0000 1.250 -0.0264 0.02556 0.01558 0.0164 1.0000 1.0000 1.500 0.0384 0.02727 0.01728 0.0072 0.9772 1.0000 1.750 0.1403 0.02920 0.01931 -0.0068 0.9273 1.0000 2.000 0.2716 0.02995 0.02035 -0.0222 0.8617 1.0000 2.250 0.4516 0.02707 0.01788 -0.0373 0.7629 1.0000 2.500 0.4847 0.02561 0.01633 -0.0331 0.6913 1.0000 2.750 0.5057 0.02471 0.01514 -0.0283 0.6321 1.0000 3.000 0.5204 0.02454 0.01465 -0.0243 0.5821 1.0000 3.250 0.5345 0.02484 0.01470 -0.0212 0.5409 1.0000 3.500 0.5514 0.02528 0.01485 -0.0188 0.5063 1.0000 3.750 0.5710 0.02590 0.01523 -0.0170 0.4768 1.0000 4.000 0.5934 0.02675 0.01587 -0.0158 0.4521 1.0000 4.250 0.6164 0.02777 0.01669 -0.0149 0.4298 1.0000 4.500 0.6389 0.02893 0.01776 -0.0142 0.4087 1.0000 4.750 0.6626 0.03026 0.01906 -0.0137 0.3909 1.0000 5.000 0.6859 0.03172 0.02047 -0.0132 0.3743 1.0000