XFOIL Version 6.94 Calculated polar for: AG35 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 0.4238 0.00689 0.00164 -0.0462 0.5706 1.0000 0.500 0.4785 0.00719 0.00179 -0.0459 0.5469 1.0000 1.000 0.5335 0.00746 0.00197 -0.0458 0.5243 1.0000 1.500 0.5888 0.00771 0.00217 -0.0457 0.5012 1.0000 2.000 0.6439 0.00796 0.00238 -0.0456 0.4753 1.0000 2.500 0.6989 0.00822 0.00259 -0.0456 0.4418 1.0000 3.000 0.7532 0.00859 0.00287 -0.0455 0.3986 1.0000 3.500 0.8063 0.00914 0.00324 -0.0453 0.3450 1.0000 4.000 0.8583 0.00986 0.00376 -0.0450 0.2873 1.0000 4.500 0.9098 0.01067 0.00439 -0.0447 0.2341 1.0000 5.000 0.9605 0.01155 0.00511 -0.0444 0.1880 1.0000 5.500 1.0108 0.01247 0.00590 -0.0441 0.1435 1.0000 6.000 1.0597 0.01354 0.00681 -0.0436 0.1027 1.0000 6.500 1.1070 0.01477 0.00793 -0.0429 0.0672 1.0000 7.000 1.1512 0.01635 0.00939 -0.0418 0.0373 1.0000 7.500 1.1924 0.01824 0.01131 -0.0402 0.0233 1.0000 8.500 1.2600 0.02298 0.01639 -0.0354 0.0153 1.0000 9.000 1.2922 0.02510 0.01875 -0.0329 0.0137 1.0000 9.500 1.3141 0.02779 0.02162 -0.0295 0.0125 1.0000 10.000 1.3118 0.03220 0.02626 -0.0236 0.0117 1.0000 10.500 1.3211 0.03585 0.03025 -0.0204 0.0113 1.0000 11.000 1.3238 0.04061 0.03536 -0.0182 0.0109 1.0000 11.500 1.3217 0.04637 0.04146 -0.0176 0.0104 1.0000 12.000 1.3152 0.05322 0.04861 -0.0186 0.0100 1.0000 12.500 1.3021 0.06176 0.05747 -0.0214 0.0097 1.0000 13.000 1.2820 0.07199 0.06803 -0.0255 0.0096 1.0000 13.500 1.2540 0.08405 0.08042 -0.0310 0.0095 1.0000 14.000 1.2200 0.09837 0.09508 -0.0382 0.0095 1.0000 14.500 1.1790 0.11559 0.11266 -0.0474 0.0097 1.0000 15.000 1.1177 0.14011 0.13760 -0.0613 0.0102 1.0000 15.500 1.0090 0.18570 0.18349 -0.0849 0.0121 1.0000