XFOIL Version 6.94 Calculated polar for: GOE 176 (ALBATROS 7020) 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 0.6440 0.00976 0.00242 -0.1311 0.5304 0.2486 0.500 0.6969 0.01005 0.00242 -0.1303 0.4798 0.2614 1.000 0.7465 0.01067 0.00264 -0.1292 0.3819 0.2937 1.500 0.7976 0.01118 0.00301 -0.1284 0.3475 0.3279 2.000 0.8506 0.01146 0.00332 -0.1280 0.3332 0.3653 2.500 0.9030 0.01099 0.00378 -0.1278 0.3229 0.7436 3.000 0.9484 0.01109 0.00392 -0.1254 0.2777 1.0000 3.500 0.9988 0.01179 0.00415 -0.1246 0.2017 1.0000 4.000 1.0349 0.01407 0.00559 -0.1217 0.0508 1.0000 4.500 1.0860 0.01461 0.00613 -0.1208 0.0460 1.0000 5.000 1.1330 0.01557 0.00690 -0.1193 0.0053 1.0000 5.500 1.1825 0.01619 0.00757 -0.1182 0.0051 1.0000 6.000 1.2306 0.01693 0.00841 -0.1168 0.0055 1.0000 6.500 1.2777 0.01770 0.00935 -0.1152 0.0061 1.0000 7.000 1.3237 0.01851 0.01026 -0.1134 0.0071 1.0000 7.500 1.3677 0.01944 0.01139 -0.1111 0.0081 1.0000 8.000 1.4081 0.02058 0.01276 -0.1084 0.0091 1.0000 8.500 1.4437 0.02194 0.01434 -0.1048 0.0110 1.0000 9.000 1.4775 0.02310 0.01567 -0.1008 0.0135 1.0000 9.500 1.4975 0.02512 0.01797 -0.0950 0.0153 1.0000 10.000 1.5123 0.02761 0.02077 -0.0891 0.0184 1.0000 10.500 1.5046 0.03222 0.02580 -0.0826 0.0206 1.0000 11.000 1.4924 0.03819 0.03206 -0.0781 0.0218 1.0000 11.500 1.4888 0.04394 0.03804 -0.0750 0.0227 1.0000 12.000 1.4890 0.04938 0.04366 -0.0702 0.0241 1.0000 12.500 1.5496 0.05287 0.04732 -0.0629 0.0258 1.0000 13.000 1.5465 0.05792 0.05258 -0.0605 0.0235 1.0000