XFOIL Version 6.96 Calculated polar for: FX 67-K-170/17 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.4985 0.04529 0.04209 -0.1262 0.7176 0.0038 -13.250 -0.5147 0.04092 0.03759 -0.1283 0.7160 0.0038 -13.000 -0.5345 0.03701 0.03352 -0.1287 0.7141 0.0038 -12.750 -0.5476 0.03416 0.03052 -0.1279 0.7124 0.0039 -12.500 -0.5703 0.03077 0.02691 -0.1257 0.7108 0.0038 -12.250 -0.5860 0.02815 0.02407 -0.1228 0.7090 0.0038 -12.000 -0.5726 0.02772 0.02364 -0.1217 0.7074 0.0040 -11.750 -0.5774 0.02611 0.02187 -0.1185 0.7056 0.0040 -11.500 -0.5799 0.02474 0.02033 -0.1150 0.7038 0.0040 -11.250 -0.5650 0.02413 0.01965 -0.1138 0.7020 0.0041 -11.000 -0.5550 0.02294 0.01830 -0.1121 0.7003 0.0043 -10.500 -0.5272 0.02079 0.01585 -0.1094 0.6972 0.0044 -10.250 -0.5105 0.01982 0.01473 -0.1083 0.6958 0.0045 -10.000 -0.4924 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0.8181 4.500 1.0390 0.00751 0.00296 -0.1128 0.5683 0.8206 5.000 1.0897 0.00774 0.00320 -0.1127 0.5554 0.8249 5.250 1.1138 0.00788 0.00335 -0.1124 0.5466 0.8269 5.500 1.1369 0.00806 0.00352 -0.1119 0.5365 0.8290 5.750 1.1557 0.00834 0.00376 -0.1106 0.5196 0.8314 6.000 1.1675 0.00880 0.00412 -0.1081 0.4942 0.8341 6.250 1.1663 0.00943 0.00462 -0.1030 0.4640 0.8373 6.500 1.1494 0.01040 0.00549 -0.0954 0.4338 0.8415 6.750 1.1354 0.01178 0.00678 -0.0893 0.4077 0.8464 7.000 1.1220 0.01342 0.00830 -0.0838 0.3779 0.8510 7.250 1.1157 0.01484 0.00964 -0.0795 0.3542 0.8551 7.500 1.1063 0.01646 0.01114 -0.0749 0.3252 0.8603 7.750 1.0951 0.01825 0.01278 -0.0703 0.2909 0.8661 8.000 1.0806 0.02027 0.01462 -0.0655 0.2516 0.8725 8.250 1.0706 0.02216 0.01635 -0.0614 0.2144 0.8797 8.500 1.0757 0.02333 0.01749 -0.0592 0.1984 0.8870 8.750 1.0669 0.02535 0.01936 -0.0557 0.1585 0.8995 9.000 1.0698 0.02690 0.02086 -0.0537 0.1310 0.9239 9.250 1.0841 0.02962 0.02338 -0.0561 0.0794 0.9564 9.500 1.0817 0.03216 0.02567 -0.0546 0.0351 1.0000 9.750 1.0980 0.03301 0.02655 -0.0542 0.0354 1.0000 10.000 1.1122 0.03401 0.02755 -0.0535 0.0337 1.0000 10.250 1.1255 0.03509 0.02866 -0.0528 0.0323 1.0000 10.500 1.1386 0.03620 0.02977 -0.0521 0.0305 1.0000 10.750 1.1531 0.03720 0.03080 -0.0516 0.0302 1.0000 11.000 1.1664 0.03831 0.03194 -0.0510 0.0288 1.0000 11.250 1.1791 0.03950 0.03316 -0.0503 0.0274 1.0000 11.500 1.1938 0.04052 0.03421 -0.0499 0.0265 1.0000 11.750 1.2080 0.04158 0.03531 -0.0494 0.0259 1.0000 12.000 1.2203 0.04282 0.03657 -0.0488 0.0227 1.0000 12.250 1.2306 0.04422 0.03796 -0.0481 0.0185 1.0000 12.500 1.2342 0.04619 0.03983 -0.0470 0.0079 1.0000 12.750 1.2390 0.04814 0.04178 -0.0459 0.0036 1.0000 13.000 1.2493 0.04961 0.04328 -0.0454 0.0031 1.0000 13.250 1.2585 0.05120 0.04491 -0.0447 0.0026 1.0000 13.500 1.2681 0.05278 0.04654 -0.0442 0.0024 1.0000 13.750 1.2780 0.05434 0.04815 -0.0437 0.0022 1.0000 14.000 1.2871 0.05600 0.04986 -0.0432 0.0021 1.0000 14.250 1.2959 0.05773 0.05163 -0.0427 0.0021 1.0000 14.500 1.3045 0.05946 0.05342 -0.0423 0.0019 1.0000 14.750 1.3105 0.06149 0.05550 -0.0418 0.0018 1.0000 15.000 1.3173 0.06346 0.05754 -0.0413 0.0017 1.0000 15.250 1.3232 0.06557 0.05972 -0.0409 0.0016 1.0000 15.500 1.3292 0.06771 0.06192 -0.0406 0.0016 1.0000 15.750 1.3361 0.06976 0.06403 -0.0403 0.0016 1.0000 16.000 1.3417 0.07198 0.06631 -0.0401 0.0015 1.0000 16.250 1.3483 0.07410 0.06849 -0.0400 0.0015 1.0000 16.500 1.3535 0.07643 0.07088 -0.0398 0.0014 1.0000 16.750 1.3586 0.07882 0.07334 -0.0398 0.0014 1.0000 17.000 1.3625 0.08135 0.07595 -0.0398 0.0014 1.0000 17.250 1.3654 0.08409 0.07877 -0.0398 0.0013 1.0000 17.500 1.3691 0.08670 0.08145 -0.0400 0.0012 1.0000 17.750 1.3738 0.08924 0.08406 -0.0403 0.0012 1.0000 18.000 1.3758 0.09216 0.08706 -0.0405 0.0011 1.0000 18.250 1.3767 0.09531 0.09030 -0.0409 0.0012 1.0000 18.500 1.3771 0.09855 0.09362 -0.0414 0.0011 1.0000 18.750 1.3791 0.10159 0.09673 -0.0420 0.0011 1.0000 19.000 1.3791 0.10495 0.10018 -0.0426 0.0010 1.0000 19.250 1.3807 0.10810 0.10340 -0.0434 0.0010 1.0000