XFOIL Version 6.96 Calculated polar for: FX 62-K-153/20 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 ------ -------- --------- --------- -------- -------- -------- -12.250 -0.4431 0.04766 0.04423 -0.1316 0.8453 0.0040 -12.000 -0.4701 0.04072 0.03698 -0.1359 0.8398 0.0039 -11.750 -0.4871 0.03570 0.03162 -0.1380 0.8350 0.0039 -11.250 -0.4932 0.02909 0.02441 -0.1395 0.8278 0.0040 -11.000 -0.4870 0.02684 0.02191 -0.1397 0.8246 0.0040 -10.750 -0.4771 0.02506 0.01987 -0.1398 0.8216 0.0041 -10.500 -0.4624 0.02332 0.01791 -0.1405 0.8189 0.0042 -10.000 -0.4211 0.02114 0.01546 -0.1419 0.8142 0.0046 -9.750 -0.3985 0.02029 0.01450 -0.1424 0.8121 0.0048 -9.500 -0.3756 0.01947 0.01355 -0.1428 0.8100 0.0050 -9.250 -0.3522 0.01867 0.01263 -0.1432 0.8080 0.0054 -9.000 -0.3285 0.01787 0.01168 -0.1435 0.8060 0.0058 -8.750 -0.3042 0.01713 0.01081 -0.1438 0.8040 0.0062 -8.500 -0.2783 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1.2933 0.01345 0.00724 -0.1593 0.3930 0.7187 6.250 1.2995 0.01439 0.00800 -0.1558 0.3500 0.7203 6.500 1.2996 0.01566 0.00899 -0.1514 0.2981 0.7219 6.750 1.3033 0.01685 0.00999 -0.1478 0.2579 0.7235 7.000 1.3103 0.01796 0.01095 -0.1450 0.2250 0.7253 7.250 1.3168 0.01916 0.01200 -0.1421 0.1910 0.7273 7.500 1.3225 0.02047 0.01313 -0.1394 0.1560 0.7293 7.750 1.3295 0.02176 0.01428 -0.1369 0.1284 0.7312 8.000 1.3375 0.02304 0.01545 -0.1346 0.1053 0.7329 8.250 1.3479 0.02417 0.01652 -0.1328 0.0888 0.7345 8.500 1.3596 0.02523 0.01756 -0.1311 0.0771 0.7361 8.750 1.3711 0.02631 0.01863 -0.1294 0.0667 0.7377 9.000 1.3825 0.02742 0.01973 -0.1277 0.0572 0.7395 9.250 1.3931 0.02859 0.02089 -0.1261 0.0477 0.7413 9.500 1.4039 0.02978 0.02207 -0.1245 0.0401 0.7431 9.750 1.4140 0.03103 0.02333 -0.1228 0.0333 0.7448 10.000 1.4254 0.03219 0.02451 -0.1214 0.0288 0.7466 10.250 1.4354 0.03347 0.02580 -0.1198 0.0247 0.7484 10.500 1.4462 0.03470 0.02708 -0.1183 0.0217 0.7501 10.750 1.4551 0.03609 0.02850 -0.1167 0.0186 0.7517 11.000 1.4657 0.03736 0.02984 -0.1153 0.0161 0.7535 11.250 1.4739 0.03885 0.03136 -0.1138 0.0137 0.7552 11.500 1.4834 0.04026 0.03285 -0.1124 0.0112 0.7569 11.750 1.4890 0.04205 0.03465 -0.1107 0.0068 0.7586 12.000 1.4938 0.04393 0.03656 -0.1091 0.0044 0.7604 12.250 1.4996 0.04576 0.03844 -0.1075 0.0037 0.7622 12.500 1.5052 0.04765 0.04041 -0.1061 0.0032 0.7640 12.750 1.5113 0.04952 0.04240 -0.1047 0.0028 0.7658 13.000 1.5176 0.05142 0.04439 -0.1035 0.0027 0.7676 13.250 1.5225 0.05349 0.04657 -0.1022 0.0025 0.7694 13.500 1.5268 0.05565 0.04884 -0.1010 0.0024 0.7713 13.750 1.5297 0.05800 0.05130 -0.0998 0.0023 0.7732 14.000 1.5324 0.06042 0.05386 -0.0987 0.0022 0.7750 14.250 1.5354 0.06287 0.05641 -0.0977 0.0021 0.7769 14.500 1.5382 0.06538 0.05904 -0.0969 0.0021 0.7787 14.750 1.5403 0.06802 0.06181 -0.0961 0.0020 0.7805 15.000 1.5414 0.07081 0.06473 -0.0953 0.0020 0.7825 15.250 1.5417 0.07375 0.06780 -0.0947 0.0020 0.7847 15.500 1.5415 0.07684 0.07102 -0.0942 0.0019 0.7870 15.750 1.5402 0.08011 0.07441 -0.0938 0.0019 0.7893 16.000 1.5375 0.08363 0.07808 -0.0936 0.0018 0.7916 16.250 1.5347 0.08724 0.08182 -0.0935 0.0018 0.7936 16.500 1.5299 0.09117 0.08590 -0.0936 0.0017 0.7956 16.750 1.5258 0.09509 0.08995 -0.0939 0.0017 0.7979 17.000 1.5205 0.09928 0.09429 -0.0944 0.0017 0.8003 17.250 1.5137 0.10376 0.09891 -0.0951 0.0016 0.8027 17.500 1.5072 0.10830 0.10361 -0.0961 0.0016 0.8052 18.000 1.4908 0.11813 0.11373 -0.0988 0.0016 0.8102 18.250 1.4825 0.12324 0.11898 -0.1005 0.0016 0.8131 18.500 1.4723 0.12880 0.12470 -0.1026 0.0015 0.8162 18.750 1.4634 0.13424 0.13029 -0.1050 0.0015 0.8197