XFOIL Version 6.96 Calculated polar for: WORTMANN FX 66-17AII-182 AIRFOIL (AS TESTED AT N 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 -0.0845 0.01055 0.00348 -0.0817 0.5290 0.0512 -4.000 -0.0321 0.00939 0.00288 -0.0822 0.5254 0.1884 -3.500 0.0198 0.00819 0.00237 -0.0827 0.5219 0.3854 -3.000 0.0762 0.00790 0.00225 -0.0833 0.5180 0.4666 -2.500 0.1337 0.00786 0.00228 -0.0839 0.5141 0.5162 -2.000 0.1925 0.00784 0.00228 -0.0847 0.5117 0.5424 -1.500 0.2514 0.00789 0.00234 -0.0854 0.5085 0.5640 -1.000 0.3105 0.00795 0.00233 -0.0863 0.5052 0.5746 -0.500 0.3691 0.00804 0.00239 -0.0871 0.5020 0.5855 0.000 0.4273 0.00822 0.00250 -0.0879 0.4977 0.5945 0.500 0.4862 0.00827 0.00257 -0.0887 0.4957 0.6021 1.000 0.5451 0.00835 0.00263 -0.0897 0.4929 0.6083 1.500 0.6036 0.00837 0.00269 -0.0905 0.4900 0.6154 2.000 0.6620 0.00847 0.00279 -0.0914 0.4872 0.6223 2.500 0.7201 0.00862 0.00291 -0.0922 0.4840 0.6279 3.000 0.7776 0.00882 0.00312 -0.0930 0.4800 0.6345 3.500 0.8358 0.00890 0.00326 -0.0939 0.4780 0.6409 4.000 0.8938 0.00901 0.00342 -0.0947 0.4754 0.6466 4.500 0.9513 0.00909 0.00358 -0.0955 0.4724 0.6531 5.000 1.0087 0.00923 0.00377 -0.0963 0.4693 0.6596 5.500 1.0655 0.00944 0.00401 -0.0970 0.4659 0.6658 6.000 1.1219 0.00960 0.00428 -0.0977 0.4616 0.6727 6.500 1.1785 0.00959 0.00435 -0.0984 0.4550 0.6799 7.000 1.2333 0.00972 0.00452 -0.0988 0.4457 0.6877 7.500 1.2895 0.00977 0.00472 -0.0994 0.4384 0.6964 8.000 1.3432 0.00999 0.00499 -0.0997 0.4287 0.7057 8.500 1.3965 0.01020 0.00527 -0.0999 0.4141 0.7161 9.000 1.4444 0.01072 0.00576 -0.0993 0.3771 0.7276 9.500 1.4521 0.01361 0.00798 -0.0929 0.2546 0.7412 10.000 1.4094 0.01853 0.01231 -0.0807 0.1392 0.7589