XFOIL Version 6.94 Calculated polar for: WACO COOTIE AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6290 0.01069 0.00321 -0.1141 0.7056 0.1019 0.500 0.6844 0.01078 0.00319 -0.1140 0.6839 0.1273 1.000 0.7395 0.01080 0.00317 -0.1139 0.6630 0.1434 1.500 0.7945 0.01088 0.00322 -0.1139 0.6436 0.1663 2.000 0.8438 0.00938 0.00343 -0.1127 0.6239 1.0000 2.500 0.8983 0.00974 0.00362 -0.1125 0.6040 1.0000 3.000 0.9523 0.01011 0.00386 -0.1123 0.5835 1.0000 3.500 1.0052 0.01043 0.00412 -0.1118 0.5570 1.0000 4.000 1.0575 0.01077 0.00443 -0.1113 0.5281 1.0000 4.500 1.1087 0.01118 0.00479 -0.1107 0.4962 1.0000 5.000 1.1582 0.01171 0.00522 -0.1098 0.4581 1.0000 5.500 1.2053 0.01241 0.00577 -0.1085 0.4141 1.0000 6.000 1.2520 0.01318 0.00646 -0.1073 0.3766 1.0000 6.500 1.2956 0.01413 0.00726 -0.1057 0.3320 1.0000 7.000 1.3356 0.01531 0.00823 -0.1035 0.2673 1.0000 7.500 1.3496 0.01859 0.01044 -0.0979 0.1006 1.0000 8.000 1.3621 0.02163 0.01298 -0.0918 0.0269 1.0000 8.500 1.3799 0.02375 0.01527 -0.0861 0.0215 1.0000 9.000 1.3949 0.02628 0.01806 -0.0810 0.0193 1.0000 9.500 1.3996 0.02985 0.02188 -0.0756 0.0180 1.0000 10.000 1.3921 0.03484 0.02710 -0.0704 0.0171 1.0000 10.500 1.3960 0.03926 0.03174 -0.0669 0.0164 1.0000 11.000 1.3981 0.04413 0.03681 -0.0638 0.0159 1.0000 11.500 1.4026 0.04897 0.04185 -0.0608 0.0156 1.0000 12.000 1.4123 0.05347 0.04655 -0.0578 0.0153 1.0000 12.500 1.4248 0.05794 0.05125 -0.0550 0.0149 1.0000 13.000 1.4358 0.06285 0.05645 -0.0525 0.0144 1.0000 13.500 1.4423 0.06884 0.06288 -0.0501 0.0149 1.0000 14.000 1.4369 0.07671 0.07124 -0.0483 0.0158 1.0000