XFOIL Version 6.96 Calculated polar for: FX 66-17AII-182 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.000 -0.6647 0.08590 0.08299 -0.0633 1.0000 0.0073 -15.750 -0.6948 0.07634 0.07323 -0.0688 1.0000 0.0072 -15.250 -0.7167 0.06012 0.05658 -0.0834 0.9341 0.0072 -15.000 -0.6539 0.05066 0.04664 -0.1081 0.8842 0.0075 -14.750 -0.6460 0.04722 0.04267 -0.1125 0.7963 0.0076 -14.500 -0.6518 0.04453 0.03971 -0.1128 0.7621 0.0076 -14.250 -0.6548 0.04216 0.03714 -0.1129 0.7375 0.0078 -14.000 -0.6572 0.03984 0.03462 -0.1129 0.7188 0.0078 -13.750 -0.6551 0.03800 0.03263 -0.1127 0.7025 0.0079 -13.500 -0.6516 0.03632 0.03081 -0.1125 0.6888 0.0081 -13.250 -0.6485 0.03460 0.02894 -0.1121 0.6773 0.0082 -13.000 -0.6456 0.03289 0.02710 -0.1116 0.6677 0.0082 -12.750 -0.6396 0.03149 0.02556 -0.1110 0.6588 0.0085 -12.500 -0.6328 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0.6358 0.01033 0.00334 -0.0829 0.4965 0.6060 2.250 0.6635 0.01039 0.00340 -0.0830 0.4947 0.6085 2.500 0.6910 0.01045 0.00345 -0.0831 0.4928 0.6109 2.750 0.7184 0.01053 0.00351 -0.0832 0.4910 0.6130 3.000 0.7456 0.01059 0.00358 -0.0832 0.4893 0.6150 3.250 0.7728 0.01068 0.00368 -0.0833 0.4876 0.6172 3.500 0.7999 0.01073 0.00379 -0.0833 0.4857 0.6195 3.750 0.8270 0.01079 0.00389 -0.0834 0.4837 0.6218 4.000 0.8541 0.01086 0.00399 -0.0834 0.4816 0.6240 4.250 0.8810 0.01093 0.00410 -0.0834 0.4795 0.6261 4.500 0.9078 0.01101 0.00420 -0.0835 0.4776 0.6283 4.750 0.9344 0.01110 0.00431 -0.0834 0.4759 0.6306 5.000 0.9608 0.01119 0.00444 -0.0834 0.4742 0.6334 5.250 0.9872 0.01130 0.00458 -0.0833 0.4726 0.6363 5.500 1.0135 0.01139 0.00475 -0.0833 0.4708 0.6393 5.750 1.0396 0.01147 0.00491 -0.0832 0.4685 0.6419 6.000 1.0649 0.01154 0.00502 -0.0829 0.4646 0.6443 6.250 1.0894 0.01160 0.00511 -0.0825 0.4603 0.6465 6.500 1.1137 0.01169 0.00525 -0.0820 0.4562 0.6488 6.750 1.1383 0.01176 0.00541 -0.0816 0.4517 0.6514 7.000 1.1623 0.01186 0.00556 -0.0812 0.4476 0.6544 7.250 1.1853 0.01198 0.00570 -0.0805 0.4435 0.6576 7.500 1.2091 0.01209 0.00590 -0.0800 0.4385 0.6607 7.750 1.2315 0.01220 0.00608 -0.0792 0.4331 0.6637 8.000 1.2523 0.01234 0.00627 -0.0782 0.4270 0.6668 8.250 1.2730 0.01248 0.00648 -0.0771 0.4187 0.6703 8.500 1.2906 0.01268 0.00670 -0.0755 0.4080 0.6740 8.750 1.2995 0.01296 0.00695 -0.0722 0.3898 0.6776 9.000 1.2983 0.01358 0.00746 -0.0673 0.3604 0.6815 9.250 1.2868 0.01468 0.00836 -0.0610 0.3220 0.6861 9.500 1.2703 0.01615 0.00965 -0.0547 0.2831 0.6912 9.750 1.2525 0.01795 0.01134 -0.0490 0.2515 0.6966 10.000 1.2309 0.02039 0.01368 -0.0440 0.2202 0.7028 10.250 1.2155 0.02297 0.01620 -0.0406 0.1965 0.7094 10.500 1.1905 0.02662 0.01975 -0.0372 0.1646 0.7169 10.750 1.1858 0.02900 0.02216 -0.0356 0.1513 0.7256 11.000 1.1698 0.03237 0.02550 -0.0336 0.1283 0.7362 11.250 1.1560 0.03569 0.02879 -0.0319 0.1072 0.7501 11.750 1.1488 0.04192 0.03522 -0.0321 0.0716 0.8390 12.000 1.1911 0.04773 0.04108 -0.0454 0.0400 0.9990 12.250 1.1919 0.05059 0.04393 -0.0454 0.0322 1.0000 12.500 1.1939 0.05310 0.04645 -0.0450 0.0275 1.0000 12.750 1.1945 0.05581 0.04916 -0.0447 0.0192 1.0000 13.000 1.1955 0.05846 0.05182 -0.0444 0.0162 1.0000 13.250 1.1976 0.06106 0.05442 -0.0442 0.0143 1.0000 13.500 1.2027 0.06336 0.05677 -0.0441 0.0136 1.0000 13.750 1.2067 0.06581 0.05928 -0.0440 0.0130 1.0000 14.000 1.2109 0.06822 0.06175 -0.0439 0.0125 1.0000 14.250 1.2129 0.07093 0.06450 -0.0438 0.0118 1.0000 14.500 1.2162 0.07349 0.06712 -0.0437 0.0114 1.0000 14.750 1.2195 0.07609 0.06979 -0.0437 0.0106 1.0000 15.000 1.2230 0.07869 0.07245 -0.0438 0.0105 1.0000 15.250 1.2270 0.08122 0.07505 -0.0438 0.0101 1.0000 15.500 1.2300 0.08389 0.07779 -0.0439 0.0098 1.0000 15.750 1.2329 0.08660 0.08056 -0.0440 0.0095 1.0000 16.000 1.2356 0.08935 0.08337 -0.0442 0.0091 1.0000 16.250 1.2363 0.09239 0.08648 -0.0445 0.0089 1.0000 16.500 1.2362 0.09551 0.08965 -0.0447 0.0086 1.0000