XFOIL Version 6.96 Calculated polar for: FX 61-184 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -12.250 -0.1812 0.09197 0.08875 -0.0913 0.9371 0.0839 -12.000 -0.5595 0.06036 0.05604 -0.1016 0.9575 0.0352 -11.750 -0.5530 0.05423 0.04972 -0.1064 0.9488 0.0343 -11.500 -0.5477 0.04852 0.04365 -0.1124 0.9393 0.0336 -11.250 -0.5402 0.04401 0.03861 -0.1185 0.9277 0.0345 -11.000 -0.5315 0.04090 0.03502 -0.1217 0.9153 0.0350 -10.750 -0.5222 0.03847 0.03216 -0.1226 0.9049 0.0353 -10.500 -0.5037 0.03555 0.02886 -0.1237 0.8980 0.0356 -10.250 -0.4787 0.03222 0.02538 -0.1241 0.8917 0.0367 -10.000 -0.4549 0.03105 0.02418 -0.1246 0.8852 0.0385 -9.750 -0.4321 0.02971 0.02267 -0.1245 0.8795 0.0397 -9.500 -0.4082 0.02834 0.02111 -0.1240 0.8738 0.0409 -9.250 -0.3782 0.02710 0.01967 -0.1242 0.8691 0.0421 -9.000 -0.3402 0.02578 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1.0671 0.01917 0.01285 -0.1071 0.5981 0.7332 5.750 1.0940 0.01909 0.01283 -0.1071 0.5916 0.7350 6.000 1.1185 0.01906 0.01287 -0.1067 0.5840 0.7371 6.250 1.1481 0.01894 0.01277 -0.1072 0.5774 0.7393 6.500 1.1716 0.01897 0.01290 -0.1067 0.5686 0.7420 6.750 1.2000 0.01892 0.01287 -0.1071 0.5603 0.7451 7.000 1.2281 0.01888 0.01284 -0.1075 0.5507 0.7478 7.250 1.2487 0.01892 0.01298 -0.1064 0.5405 0.7498 7.500 1.2745 0.01889 0.01297 -0.1061 0.5310 0.7519 7.750 1.2945 0.01899 0.01315 -0.1049 0.5197 0.7543 8.000 1.3143 0.01915 0.01339 -0.1038 0.5082 0.7570 8.250 1.3345 0.01929 0.01357 -0.1028 0.4954 0.7600 8.500 1.3532 0.01948 0.01376 -0.1016 0.4810 0.7633 8.750 1.3662 0.01969 0.01401 -0.0992 0.4660 0.7658 9.000 1.3745 0.01996 0.01432 -0.0961 0.4506 0.7685 9.250 1.3805 0.02038 0.01476 -0.0927 0.4336 0.7714 9.500 1.3859 0.02097 0.01536 -0.0896 0.4155 0.7745 9.750 1.3909 0.02173 0.01609 -0.0867 0.3969 0.7778 10.000 1.3936 0.02264 0.01698 -0.0838 0.3784 0.7808 10.250 1.3938 0.02373 0.01806 -0.0806 0.3596 0.7840 10.500 1.3938 0.02503 0.01936 -0.0779 0.3398 0.7878 10.750 1.3934 0.02656 0.02085 -0.0755 0.3195 0.7917 11.000 1.3923 0.02832 0.02253 -0.0734 0.2998 0.7954 11.250 1.3905 0.03010 0.02430 -0.0711 0.2810 0.7986 11.500 1.3892 0.03202 0.02622 -0.0692 0.2623 0.8023 11.750 1.3879 0.03409 0.02825 -0.0676 0.2445 0.8064 12.000 1.3868 0.03632 0.03044 -0.0662 0.2275 0.8106 12.250 1.3846 0.03857 0.03268 -0.0647 0.2113 0.8145 12.500 1.3825 0.04094 0.03506 -0.0634 0.1957 0.8193 12.750 1.3805 0.04353 0.03762 -0.0625 0.1804 0.8245 13.000 1.3777 0.04621 0.04030 -0.0615 0.1652 0.8295 13.250 1.3741 0.04910 0.04319 -0.0607 0.1497 0.8350 13.500 1.3704 0.05223 0.04630 -0.0602 0.1336 0.8411 13.750 1.3648 0.05551 0.04958 -0.0596 0.1185 0.8471 14.000 1.3584 0.05907 0.05311 -0.0592 0.1051 0.8541 14.250 1.3506 0.06285 0.05685 -0.0589 0.0944 0.8618 14.500 1.3470 0.06621 0.06026 -0.0586 0.0849 0.8722 14.750 1.3422 0.06948 0.06361 -0.0581 0.0779 0.8870 15.000 1.3313 0.07187 0.06613 -0.0559 0.0736 0.9745 15.250 1.3310 0.07568 0.06992 -0.0569 0.0686 1.0000 15.500 1.3350 0.07902 0.07333 -0.0579 0.0640 1.0000 15.750 1.3352 0.08265 0.07691 -0.0588 0.0604 1.0000 16.000 1.3400 0.08575 0.08008 -0.0596 0.0572 1.0000 16.250 1.3439 0.08902 0.08343 -0.0606 0.0541 1.0000 16.500 1.3461 0.09246 0.08690 -0.0617 0.0514 1.0000 16.750 1.3504 0.09537 0.08978 -0.0622 0.0487 1.0000 17.000 1.3540 0.09873 0.09330 -0.0634 0.0466 1.0000 17.250 1.3567 0.10216 0.09681 -0.0646 0.0444 1.0000 17.500 1.3584 0.10570 0.10039 -0.0660 0.0424 1.0000 17.750 1.3634 0.10843 0.10308 -0.0666 0.0402 1.0000 18.000 1.3640 0.11237 0.10721 -0.0683 0.0386 1.0000 18.250 1.3658 0.11599 0.11095 -0.0698 0.0370 1.0000 18.500 1.3678 0.11952 0.11454 -0.0714 0.0355 1.0000 18.750 1.3708 0.12273 0.11776 -0.0728 0.0339 1.0000 19.000 1.3742 0.12583 0.12095 -0.0739 0.0323 1.0000 19.250 1.3717 0.13035 0.12566 -0.0762 0.0309 1.0000