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 = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.500 -0.5417 0.07651 0.07269 -0.0755 1.0000 0.0264 -12.250 -0.5809 0.06887 0.06496 -0.0788 1.0000 0.0263 -12.000 -0.6130 0.06208 0.05800 -0.0825 0.9986 0.0262 -11.750 -0.6444 0.05320 0.04868 -0.0906 0.9914 0.0265 -11.500 -0.6512 0.04841 0.04361 -0.0952 0.9848 0.0269 -11.250 -0.6432 0.04606 0.04116 -0.0976 0.9779 0.0273 -11.000 -0.6305 0.04381 0.03877 -0.1006 0.9722 0.0277 -10.750 -0.6196 0.04232 0.03716 -0.1015 0.9633 0.0283 -10.500 -0.6004 0.03992 0.03455 -0.1042 0.9589 0.0291 -10.250 -0.5919 0.03763 0.03200 -0.1041 0.9494 0.0297 -10.000 -0.5715 0.03492 0.02894 -0.1061 0.9451 0.0308 -9.750 -0.5594 0.03287 0.02655 -0.1055 0.9351 0.0318 -9.500 -0.5358 0.03085 0.02409 -0.1068 0.9302 0.0328 -9.250 -0.5193 0.02912 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-0.0395 0.0545 0.8429 10.750 1.1053 0.03670 0.02839 -0.0387 0.0534 0.8601 11.000 1.1182 0.03818 0.03009 -0.0389 0.0523 0.9363 11.250 1.1227 0.03971 0.03168 -0.0373 0.0514 1.0000 11.500 1.1310 0.04138 0.03340 -0.0364 0.0504 1.0000 11.750 1.1394 0.04306 0.03514 -0.0357 0.0494 1.0000 12.000 1.1475 0.04479 0.03691 -0.0350 0.0483 1.0000 12.250 1.1546 0.04663 0.03876 -0.0343 0.0470 1.0000 12.500 1.1623 0.04844 0.04060 -0.0336 0.0458 1.0000 12.750 1.1734 0.04997 0.04224 -0.0332 0.0445 1.0000 13.000 1.1836 0.05161 0.04397 -0.0328 0.0431 1.0000 13.250 1.1931 0.05334 0.04578 -0.0324 0.0416 1.0000 13.500 1.2020 0.05514 0.04766 -0.0321 0.0405 1.0000 13.750 1.2100 0.05706 0.04964 -0.0319 0.0391 1.0000 14.000 1.2167 0.05915 0.05176 -0.0316 0.0379 1.0000 14.250 1.2243 0.06115 0.05381 -0.0313 0.0368 1.0000 14.500 1.2332 0.06308 0.05589 -0.0312 0.0357 1.0000 14.750 1.2413 0.06513 0.05807 -0.0312 0.0343 1.0000 15.000 1.2482 0.06735 0.06039 -0.0313 0.0328 1.0000 15.250 1.2542 0.06972 0.06286 -0.0315 0.0316 1.0000 15.500 1.2587 0.07232 0.06552 -0.0319 0.0304 1.0000 15.750 1.2626 0.07499 0.06825 -0.0321 0.0292 1.0000 16.000 1.2694 0.07738 0.07083 -0.0324 0.0278 1.0000 16.250 1.2745 0.08002 0.07361 -0.0328 0.0263 1.0000 16.500 1.2780 0.08293 0.07662 -0.0335 0.0249 1.0000 16.750 1.2798 0.08615 0.07990 -0.0343 0.0238 1.0000 17.000 1.2834 0.08908 0.08300 -0.0349 0.0222 1.0000 17.250 1.2864 0.09219 0.08625 -0.0357 0.0204 1.0000 17.500 1.2879 0.09557 0.08974 -0.0368 0.0193 1.0000 17.750 1.2888 0.09904 0.09331 -0.0379 0.0182 1.0000 18.000 1.2896 0.10252 0.09694 -0.0389 0.0168 1.0000 18.250 1.2897 0.10621 0.10074 -0.0402 0.0158 1.0000 18.500 1.2888 0.11012 0.10474 -0.0419 0.0151 1.0000 18.750 1.2866 0.11425 0.10898 -0.0435 0.0145 1.0000 19.000 1.2839 0.11850 0.11337 -0.0452 0.0139 1.0000 19.250 1.2802 0.12295 0.11796 -0.0472 0.0135 1.0000