XFOIL Version 6.96 Calculated polar for: FX 63-158 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 ------ -------- --------- --------- -------- -------- -------- -11.000 -0.1866 0.09893 0.09686 -0.0606 0.9902 0.0174 -10.750 -0.1867 0.09237 0.09031 -0.0648 0.9876 0.0178 -10.500 -0.3018 0.10308 0.10094 -0.0490 0.9985 0.0161 -10.250 -0.1988 0.07488 0.07284 -0.0757 0.9802 0.0180 -10.000 -0.2119 0.06618 0.06409 -0.0812 0.9756 0.0180 -9.750 -0.2283 0.05909 0.05693 -0.0859 0.9613 0.0180 -9.500 -0.2444 0.05275 0.05048 -0.0906 0.9491 0.0180 -9.250 -0.2561 0.04682 0.04440 -0.0962 0.9368 0.0180 -9.000 -0.2661 0.04080 0.03809 -0.1043 0.9202 0.0180 -8.750 -0.3000 0.05407 0.05161 -0.0973 0.9327 0.0175 -8.500 -0.3014 0.04622 0.04297 -0.1107 0.9096 0.0180 -8.250 -0.2313 0.03782 0.03417 -0.1281 0.8751 0.0185 -8.000 -0.2027 0.03523 0.03114 -0.1319 0.8056 0.0188 -7.750 -0.1923 0.03352 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0.6852 2.500 0.8910 0.01165 0.00428 -0.1335 0.3362 0.6896 2.750 0.9154 0.01190 0.00454 -0.1329 0.3233 0.6943 3.000 0.9381 0.01237 0.00488 -0.1320 0.2949 0.7000 3.250 0.9637 0.01271 0.00513 -0.1317 0.2855 0.7049 3.500 0.9888 0.01287 0.00533 -0.1313 0.2778 0.7078 3.750 1.0117 0.01326 0.00564 -0.1306 0.2477 0.7100 4.000 1.0334 0.01374 0.00599 -0.1297 0.2299 0.7126 4.250 1.0579 0.01402 0.00626 -0.1292 0.2209 0.7153 4.500 1.0823 0.01432 0.00653 -0.1289 0.1964 0.7177 4.750 1.1042 0.01481 0.00689 -0.1281 0.1817 0.7196 5.000 1.1289 0.01508 0.00715 -0.1278 0.1775 0.7215 5.250 1.1531 0.01534 0.00741 -0.1274 0.1738 0.7234 5.500 1.1768 0.01554 0.00767 -0.1268 0.1712 0.7256 5.750 1.2006 0.01574 0.00793 -0.1262 0.1686 0.7282 6.000 1.2239 0.01599 0.00823 -0.1256 0.1635 0.7311 6.250 1.2470 0.01625 0.00854 -0.1249 0.1477 0.7343 6.500 1.2647 0.01686 0.00904 -0.1235 0.1360 0.7369 6.750 1.2854 0.01727 0.00943 -0.1225 0.1321 0.7390 7.000 1.3037 0.01765 0.00983 -0.1211 0.1280 0.7410 7.250 1.3225 0.01794 0.01019 -0.1197 0.1248 0.7430 7.500 1.3411 0.01826 0.01058 -0.1183 0.1197 0.7449 7.750 1.3590 0.01864 0.01102 -0.1168 0.1148 0.7468 8.000 1.3718 0.01931 0.01163 -0.1147 0.0965 0.7485 8.250 1.3842 0.02000 0.01229 -0.1125 0.0925 0.7501 8.500 1.3970 0.02069 0.01300 -0.1105 0.0884 0.7515 8.750 1.4122 0.02126 0.01365 -0.1089 0.0850 0.7528 9.000 1.4232 0.02207 0.01451 -0.1068 0.0813 0.7540 9.250 1.4390 0.02264 0.01515 -0.1054 0.0780 0.7552 9.500 1.4517 0.02341 0.01591 -0.1037 0.0643 0.7562 9.750 1.4582 0.02453 0.01699 -0.1014 0.0618 0.7576 10.000 1.4649 0.02570 0.01820 -0.0993 0.0595 0.7589 10.250 1.4717 0.02695 0.01952 -0.0974 0.0576 0.7602 10.500 1.4772 0.02837 0.02106 -0.0955 0.0562 0.7617 10.750 1.4793 0.03016 0.02297 -0.0937 0.0545 0.7633 11.000 1.4775 0.03242 0.02532 -0.0918 0.0528 0.7650 11.250 1.4734 0.03509 0.02808 -0.0902 0.0510 0.7669 11.500 1.4889 0.03610 0.02920 -0.0899 0.0491 0.7687 11.750 1.4988 0.03769 0.03089 -0.0894 0.0463 0.7703 12.000 1.5053 0.03967 0.03293 -0.0888 0.0436 0.7717 12.250 1.5216 0.04072 0.03408 -0.0887 0.0390 0.7732 12.500 1.5288 0.04269 0.03606 -0.0884 0.0347 0.7747 12.750 1.5319 0.04514 0.03850 -0.0880 0.0318 0.7763 13.000 1.5344 0.04772 0.04115 -0.0877 0.0297 0.7778 13.250 1.5362 0.05044 0.04397 -0.0875 0.0282 0.7794 13.500 1.5366 0.05340 0.04701 -0.0874 0.0267 0.7810 13.750 1.5362 0.05653 0.05023 -0.0873 0.0253 0.7827 14.000 1.5353 0.05981 0.05359 -0.0874 0.0239 0.7844 14.250 1.5393 0.06256 0.05644 -0.0877 0.0231 0.7863 14.500 1.5432 0.06537 0.05934 -0.0880 0.0215 0.7883 14.750 1.5446 0.06859 0.06264 -0.0884 0.0194 0.7902 15.000 1.5447 0.07203 0.06614 -0.0889 0.0174 0.7921 15.250 1.5450 0.07548 0.06966 -0.0895 0.0154 0.7941 15.500 1.5433 0.07931 0.07354 -0.0902 0.0135 0.7959 15.750 1.5431 0.08300 0.07732 -0.0911 0.0120 0.7979 16.000 1.5414 0.08697 0.08136 -0.0920 0.0104 0.7999 16.250 1.5413 0.09080 0.08529 -0.0930 0.0092 0.8020 16.500 1.5390 0.09498 0.08955 -0.0942 0.0067 0.8041 16.750 1.5340 0.09968 0.09431 -0.0956 0.0049 0.8061 17.000 1.5306 0.10421 0.09892 -0.0971 0.0046 0.8081