XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0192 0.02940 0.01461 -0.0239 0.9994 1.0006 -2.750 0.0019 0.02978 0.01475 -0.0236 0.9994 1.0006 -2.500 0.0229 0.03023 0.01496 -0.0232 0.9994 1.0006 -2.250 0.0437 0.03075 0.01528 -0.0228 0.9994 1.0006 -2.000 0.0644 0.03133 0.01568 -0.0224 0.9994 1.0006 -1.750 0.0852 0.03196 0.01616 -0.0220 0.9994 1.0006 -1.500 0.1059 0.03265 0.01671 -0.0218 0.9994 1.0006 -1.250 0.1267 0.03339 0.01734 -0.0215 0.9994 1.0006 -1.000 0.1474 0.03418 0.01804 -0.0214 0.9994 1.0006 -0.750 0.1680 0.03502 0.01882 -0.0212 0.9994 1.0006 -0.500 0.1884 0.03591 0.01967 -0.0212 0.9994 1.0006 -0.250 0.2086 0.03686 0.02061 -0.0212 0.9994 1.0006 0.000 0.2285 0.03788 0.02163 -0.0212 0.9994 1.0006 0.250 0.2481 0.03898 0.02276 -0.0213 0.9994 1.0006 0.500 0.2672 0.04017 0.02401 -0.0215 0.9994 1.0006 0.750 0.2856 0.04148 0.02541 -0.0218 0.9994 1.0006 1.000 0.3030 0.04296 0.02701 -0.0223 0.9994 1.0006 1.250 0.3188 0.04470 0.02890 -0.0229 0.9994 1.0006 1.500 0.3317 0.04689 0.03129 -0.0240 0.9994 1.0006 1.750 0.3384 0.05006 0.03468 -0.0258 0.9994 1.0006 2.000 0.6119 0.04966 0.03528 -0.0716 0.8156 1.0006 2.250 0.6677 0.04968 0.03547 -0.0723 0.7390 1.0006 2.500 0.7032 0.05020 0.03601 -0.0702 0.6872 1.0006 2.750 0.7345 0.05088 0.03665 -0.0677 0.6480 1.0006 3.000 0.7591 0.05223 0.03805 -0.0657 0.6140 1.0006 3.250 0.7818 0.05391 0.03976 -0.0641 0.5840 1.0006 3.500 0.8032 0.05663 0.04257 -0.0641 0.5611 1.0006 3.750 0.8230 0.05932 0.04534 -0.0638 0.5388 1.0006 4.000 0.8466 0.06129 0.04738 -0.0625 0.5175 1.0006 4.250 0.8568 0.06598 0.05225 -0.0640 0.5043 1.0006 4.500 0.8664 0.07042 0.05680 -0.0647 0.4920 1.0006 4.750 0.8788 0.07416 0.06062 -0.0646 0.4783 1.0006 5.000 0.8723 0.08030 0.06686 -0.0660 0.4711 1.0006