XFOIL Version 6.96 Calculated polar for: FX 60-160 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.000 -0.5292 0.08378 0.08138 -0.0698 1.0000 0.0247 -12.750 -0.5686 0.07431 0.07185 -0.0744 1.0000 0.0244 -12.500 -0.5963 0.06851 0.06600 -0.0766 1.0000 0.0240 -12.250 -0.6337 0.06150 0.05886 -0.0799 0.9995 0.0239 -12.000 -0.6700 0.05099 0.04798 -0.0899 0.9951 0.0238 -11.750 -0.7018 0.04242 0.03887 -0.0969 0.9878 0.0242 -11.250 -0.6980 0.03420 0.02996 -0.1033 0.9742 0.0254 -11.000 -0.6685 0.03319 0.02892 -0.1060 0.9717 0.0260 -10.750 -0.6362 0.03234 0.02803 -0.1089 0.9700 0.0266 -10.500 -0.6161 0.03121 0.02680 -0.1094 0.9636 0.0274 -10.250 -0.5915 0.02909 0.02440 -0.1113 0.9598 0.0286 -10.000 -0.5618 0.02732 0.02227 -0.1137 0.9574 0.0297 -9.750 -0.5335 0.02436 0.01906 -0.1161 0.9554 0.0307 -9.500 -0.5126 0.02366 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5.000 0.9735 0.01008 0.00454 -0.1031 0.5441 0.7347 5.250 0.9933 0.01025 0.00465 -0.1016 0.5244 0.7374 5.500 1.0106 0.01045 0.00479 -0.0996 0.5022 0.7401 5.750 1.0274 0.01069 0.00498 -0.0975 0.4782 0.7428 6.000 1.0414 0.01101 0.00521 -0.0949 0.4543 0.7456 6.250 1.0537 0.01133 0.00544 -0.0920 0.4282 0.7487 6.500 1.0632 0.01179 0.00577 -0.0887 0.3980 0.7518 6.750 1.0689 0.01240 0.00620 -0.0847 0.3592 0.7546 7.000 1.0672 0.01335 0.00687 -0.0796 0.2998 0.7574 7.250 1.0509 0.01503 0.00799 -0.0725 0.2012 0.7607 7.500 1.0317 0.01704 0.00943 -0.0655 0.1020 0.7643 7.750 1.0351 0.01812 0.01038 -0.0621 0.0788 0.7675 8.000 1.0438 0.01894 0.01121 -0.0595 0.0701 0.7705 8.250 1.0561 0.01961 0.01194 -0.0575 0.0657 0.7736 8.500 1.0653 0.02048 0.01283 -0.0551 0.0618 0.7773 8.750 1.0738 0.02144 0.01384 -0.0528 0.0588 0.7810 9.000 1.0866 0.02218 0.01465 -0.0512 0.0571 0.7843 9.250 1.0980 0.02300 0.01555 -0.0494 0.0554 0.7880 9.500 1.1087 0.02391 0.01652 -0.0476 0.0541 0.7922 9.750 1.1185 0.02492 0.01759 -0.0459 0.0527 0.7967 10.000 1.1250 0.02618 0.01889 -0.0439 0.0507 0.8010 10.250 1.1257 0.02786 0.02066 -0.0415 0.0491 0.8058 10.500 1.1385 0.02879 0.02167 -0.0404 0.0485 0.8114 10.750 1.1500 0.02981 0.02278 -0.0392 0.0476 0.8176 11.000 1.1613 0.03087 0.02394 -0.0380 0.0465 0.8254 11.250 1.1720 0.03200 0.02517 -0.0369 0.0454 0.8349 11.500 1.1817 0.03321 0.02650 -0.0357 0.0444 0.8498 11.750 1.2010 0.03446 0.02795 -0.0366 0.0431 0.8947 12.000 1.2171 0.03630 0.02988 -0.0379 0.0418 1.0000 12.250 1.2185 0.03849 0.03210 -0.0366 0.0407 1.0000 12.500 1.2299 0.03987 0.03354 -0.0360 0.0400 1.0000 12.750 1.2424 0.04117 0.03491 -0.0355 0.0391 1.0000 13.000 1.2534 0.04262 0.03642 -0.0350 0.0382 1.0000 13.250 1.2637 0.04414 0.03800 -0.0345 0.0373 1.0000 13.500 1.2742 0.04568 0.03958 -0.0341 0.0362 1.0000 13.750 1.2838 0.04732 0.04125 -0.0337 0.0352 1.0000 14.000 1.2889 0.04939 0.04334 -0.0331 0.0340 1.0000 14.250 1.2946 0.05144 0.04545 -0.0326 0.0329 1.0000 14.500 1.3070 0.05295 0.04705 -0.0326 0.0317 1.0000 14.750 1.3180 0.05460 0.04877 -0.0326 0.0303 1.0000 15.000 1.3282 0.05635 0.05055 -0.0326 0.0289 1.0000 15.250 1.3313 0.05886 0.05307 -0.0324 0.0275 1.0000 15.500 1.3426 0.06055 0.05487 -0.0325 0.0257 1.0000 15.750 1.3539 0.06229 0.05665 -0.0327 0.0236 1.0000 16.000 1.3597 0.06465 0.05904 -0.0328 0.0215 1.0000 16.250 1.3680 0.06673 0.06119 -0.0329 0.0195 1.0000 16.500 1.3709 0.06950 0.06396 -0.0332 0.0181 1.0000 16.750 1.3746 0.07217 0.06673 -0.0333 0.0169 1.0000 17.000 1.3786 0.07488 0.06950 -0.0337 0.0158 1.0000 17.250 1.3798 0.07799 0.07266 -0.0341 0.0150 1.0000 17.500 1.3779 0.08157 0.07629 -0.0347 0.0144 1.0000 17.750 1.3791 0.08475 0.07958 -0.0353 0.0139 1.0000 18.000 1.3782 0.08828 0.08321 -0.0360 0.0134 1.0000 18.250 1.3788 0.09165 0.08666 -0.0369 0.0128 1.0000 18.500 1.3770 0.09542 0.09052 -0.0379 0.0125 1.0000 18.750 1.3744 0.09939 0.09455 -0.0391 0.0121 1.0000 19.000 1.3685 0.10387 0.09911 -0.0405 0.0118 1.0000 19.250 1.3629 0.10836 0.10369 -0.0421 0.0116 1.0000