XFOIL Version 6.96 Calculated polar for: NASA NLF1015 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 ------ -------- --------- --------- -------- -------- -------- -11.250 -0.0924 0.08680 0.08494 -0.1339 0.9421 0.0141 -11.000 -0.0796 0.08018 0.07829 -0.1415 0.9381 0.0152 -10.750 -0.1041 0.05465 0.05257 -0.1672 0.9292 0.0164 -10.500 -0.1027 0.04378 0.04131 -0.1876 0.9132 0.0164 -10.250 -0.1089 0.03946 0.03668 -0.1929 0.8914 0.0164 -10.000 -0.1267 0.03694 0.03392 -0.1920 0.8722 0.0164 -9.750 -0.1440 0.03532 0.03212 -0.1888 0.8578 0.0164 -9.500 -0.1532 0.03335 0.02991 -0.1866 0.8473 0.0164 -8.750 -0.1803 0.01985 0.01507 -0.1784 0.8235 0.0127 -8.500 -0.1615 0.01866 0.01369 -0.1778 0.8188 0.0127 -8.250 -0.1410 0.01778 0.01269 -0.1772 0.8146 0.0126 -8.000 -0.1216 0.01633 0.01108 -0.1765 0.8103 0.0128 -7.750 -0.1005 0.01511 0.00973 -0.1760 0.8064 0.0130 -7.500 -0.0784 0.01417 0.00870 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0.2896 0.7751 7.500 1.3243 0.01529 0.00946 -0.1429 0.2479 0.7809 7.750 1.3265 0.01650 0.01049 -0.1392 0.2113 0.7874 8.000 1.3315 0.01764 0.01150 -0.1360 0.1799 0.7936 8.250 1.3363 0.01886 0.01258 -0.1329 0.1493 0.8005 8.500 1.3406 0.02018 0.01375 -0.1299 0.1186 0.8075 8.750 1.3442 0.02159 0.01503 -0.1269 0.0892 0.8153 9.000 1.3466 0.02317 0.01644 -0.1239 0.0582 0.8233 9.250 1.3478 0.02491 0.01802 -0.1209 0.0298 0.8324 9.500 1.3507 0.02659 0.01962 -0.1181 0.0123 0.8421 9.750 1.3633 0.02761 0.02070 -0.1166 0.0104 0.8535 10.000 1.3753 0.02868 0.02183 -0.1150 0.0094 0.8670 10.250 1.3858 0.02983 0.02310 -0.1132 0.0086 0.8838 10.500 1.3966 0.03078 0.02416 -0.1114 0.0083 0.9110 10.750 1.4033 0.03165 0.02517 -0.1088 0.0080 1.0000 11.000 1.4158 0.03288 0.02645 -0.1076 0.0076 1.0000 11.250 1.4278 0.03416 0.02778 -0.1065 0.0073 1.0000 11.500 1.4383 0.03561 0.02927 -0.1052 0.0070 1.0000 11.750 1.4478 0.03715 0.03087 -0.1038 0.0068 1.0000 12.000 1.4553 0.03891 0.03271 -0.1024 0.0066 1.0000 12.250 1.4601 0.04098 0.03487 -0.1008 0.0064 1.0000 12.500 1.4647 0.04310 0.03708 -0.0993 0.0062 1.0000 12.750 1.4737 0.04482 0.03887 -0.0982 0.0062 1.0000 13.000 1.4819 0.04666 0.04077 -0.0971 0.0061 1.0000 13.250 1.4892 0.04862 0.04280 -0.0961 0.0060 1.0000 13.500 1.4957 0.05070 0.04496 -0.0951 0.0059 1.0000 13.750 1.5014 0.05293 0.04727 -0.0941 0.0059 1.0000 14.000 1.5068 0.05521 0.04964 -0.0932 0.0058 1.0000 14.250 1.5107 0.05772 0.05223 -0.0923 0.0057 1.0000 14.500 1.5149 0.06023 0.05482 -0.0915 0.0056 1.0000 14.750 1.5184 0.06288 0.05756 -0.0908 0.0055 1.0000 15.000 1.5212 0.06570 0.06046 -0.0902 0.0054 1.0000 15.250 1.5239 0.06856 0.06341 -0.0897 0.0053 1.0000 15.500 1.5257 0.07157 0.06651 -0.0893 0.0053 1.0000 15.750 1.5268 0.07475 0.06978 -0.0890 0.0052 1.0000 16.000 1.5277 0.07804 0.07316 -0.0888 0.0051 1.0000 16.250 1.5283 0.08141 0.07662 -0.0887 0.0051 1.0000 16.500 1.5281 0.08496 0.08027 -0.0888 0.0050 1.0000 16.750 1.5274 0.08862 0.08404 -0.0890 0.0049 1.0000 17.000 1.5262 0.09241 0.08792 -0.0893 0.0049 1.0000 17.250 1.5248 0.09629 0.09189 -0.0898 0.0048 1.0000 17.500 1.5229 0.10029 0.09598 -0.0905 0.0048 1.0000 17.750 1.5207 0.10439 0.10018 -0.0913 0.0047 1.0000 18.000 1.5178 0.10860 0.10449 -0.0923 0.0046 1.0000 18.250 1.5133 0.11303 0.10902 -0.0933 0.0046 1.0000 18.500 1.5082 0.11755 0.11366 -0.0944 0.0045 1.0000 18.750 1.5029 0.12215 0.11838 -0.0957 0.0044 1.0000 19.000 1.4971 0.12693 0.12330 -0.0973 0.0044 1.0000 19.250 1.4926 0.13162 0.12812 -0.0993 0.0044 1.0000