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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6280 0.00933 0.00221 -0.1133 0.6631 0.0703 0.500 0.6836 0.00952 0.00235 -0.1132 0.6434 0.1035 1.000 0.7394 0.00966 0.00245 -0.1133 0.6253 0.1186 1.500 0.7947 0.00980 0.00256 -0.1133 0.6080 0.1338 2.000 0.8503 0.00985 0.00268 -0.1135 0.5901 0.1570 2.500 0.8994 0.00841 0.00294 -0.1125 0.5723 1.0000 3.000 0.9537 0.00869 0.00310 -0.1123 0.5478 1.0000 3.500 1.0076 0.00898 0.00333 -0.1122 0.5219 1.0000 4.000 1.0605 0.00934 0.00361 -0.1118 0.4900 1.0000 4.500 1.1115 0.00985 0.00398 -0.1113 0.4497 1.0000 5.000 1.1606 0.01054 0.00446 -0.1104 0.4034 1.0000 5.500 1.2094 0.01125 0.00504 -0.1096 0.3641 1.0000 6.000 1.2565 0.01207 0.00571 -0.1086 0.3238 1.0000 6.500 1.3008 0.01310 0.00650 -0.1071 0.2646 1.0000 7.000 1.3191 0.01639 0.00864 -0.1021 0.0861 1.0000 7.500 1.3437 0.01884 0.01070 -0.0976 0.0209 1.0000 8.000 1.3735 0.02055 0.01262 -0.0937 0.0167 1.0000 8.500 1.3954 0.02235 0.01462 -0.0886 0.0153 1.0000 9.000 1.4090 0.02483 0.01729 -0.0830 0.0141 1.0000 9.500 1.4088 0.02862 0.02130 -0.0770 0.0129 1.0000 10.000 1.4000 0.03366 0.02654 -0.0714 0.0124 1.0000 10.500 1.4105 0.03741 0.03049 -0.0683 0.0120 1.0000 11.000 1.4165 0.04181 0.03508 -0.0651 0.0116 1.0000 11.500 1.4238 0.04627 0.03973 -0.0623 0.0111 1.0000 12.000 1.4319 0.05079 0.04443 -0.0599 0.0105 1.0000 12.500 1.4416 0.05531 0.04912 -0.0573 0.0102 1.0000 13.000 1.4527 0.05995 0.05397 -0.0546 0.0101 1.0000 13.500 1.4567 0.06587 0.06029 -0.0522 0.0103 1.0000 14.000 1.4497 0.07388 0.06881 -0.0499 0.0113 1.0000