XFOIL Version 6.96 Calculated polar for: FX 61-168 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.3838 0.09513 0.09029 -0.0582 1.0000 0.0222 -11.000 -0.4120 0.08400 0.07915 -0.0646 1.0000 0.0213 -10.750 -0.4696 0.07229 0.06725 -0.0710 1.0000 0.0202 -10.250 -0.5254 0.06402 0.05870 -0.0708 1.0000 0.0197 -10.000 -0.5473 0.06157 0.05621 -0.0685 1.0000 0.0197 -9.750 -0.5516 0.05745 0.05189 -0.0714 0.9901 0.0196 -9.500 -0.5529 0.05311 0.04722 -0.0745 0.9678 0.0197 -9.250 -0.5456 0.04904 0.04278 -0.0768 0.9464 0.0197 -9.000 -0.5342 0.04524 0.03861 -0.0784 0.9280 0.0199 -8.750 -0.5130 0.04236 0.03555 -0.0806 0.9138 0.0206 -8.500 -0.4855 0.04007 0.03305 -0.0832 0.9012 0.0218 -8.250 -0.4550 0.03725 0.02982 -0.0860 0.8894 0.0230 -8.000 -0.4228 0.03439 0.02649 -0.0883 0.8763 0.0240 -7.750 -0.3872 0.03186 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1.1041 0.02442 0.01638 -0.0848 0.5029 0.7991 7.250 1.1458 0.02474 0.01698 -0.0831 0.4855 0.8051 7.500 1.1697 0.02471 0.01701 -0.0824 0.4774 0.8082 7.750 1.1850 0.02504 0.01751 -0.0808 0.4667 0.8117 8.000 1.2053 0.02526 0.01785 -0.0799 0.4568 0.8151 8.250 1.2280 0.02538 0.01804 -0.0793 0.4476 0.8182 8.500 1.2385 0.02576 0.01859 -0.0768 0.4365 0.8215 8.750 1.2518 0.02608 0.01902 -0.0747 0.4254 0.8253 9.000 1.2681 0.02642 0.01942 -0.0732 0.4141 0.8292 9.250 1.2851 0.02680 0.01988 -0.0719 0.4026 0.8329 9.500 1.2934 0.02745 0.02067 -0.0694 0.3905 0.8364 9.750 1.3020 0.02816 0.02148 -0.0672 0.3774 0.8406 10.000 1.3103 0.02900 0.02238 -0.0651 0.3635 0.8454 10.250 1.3154 0.02995 0.02342 -0.0628 0.3493 0.8499 10.500 1.3183 0.03110 0.02463 -0.0604 0.3348 0.8548 10.750 1.3199 0.03253 0.02612 -0.0583 0.3188 0.8600 11.000 1.3190 0.03414 0.02778 -0.0562 0.3027 0.8651 11.250 1.3166 0.03598 0.02967 -0.0542 0.2863 0.8711 11.500 1.3133 0.03811 0.03181 -0.0526 0.2694 0.8778 11.750 1.3079 0.04041 0.03417 -0.0510 0.2527 0.8856 12.000 1.3030 0.04283 0.03663 -0.0497 0.2369 0.8949 12.500 1.2906 0.04781 0.04169 -0.0469 0.2090 0.9270 12.750 1.2834 0.05050 0.04444 -0.0461 0.1956 1.0000 13.000 1.2816 0.05376 0.04769 -0.0465 0.1816 1.0000 13.500 1.2773 0.06059 0.05456 -0.0477 0.1558 1.0000 13.750 1.2748 0.06420 0.05821 -0.0484 0.1435 1.0000 14.000 1.2717 0.06797 0.06201 -0.0493 0.1320 1.0000 14.250 1.2681 0.07188 0.06596 -0.0502 0.1213 1.0000 14.500 1.2636 0.07602 0.07016 -0.0513 0.1111 1.0000 14.750 1.2581 0.08040 0.07456 -0.0526 0.1013 1.0000 15.000 1.2537 0.08475 0.07896 -0.0540 0.0916 1.0000 15.250 1.2488 0.08927 0.08352 -0.0555 0.0831 1.0000 15.500 1.2421 0.09412 0.08837 -0.0572 0.0760 1.0000 15.750 1.2392 0.09853 0.09288 -0.0589 0.0688 1.0000 16.000 1.2326 0.10354 0.09789 -0.0609 0.0635 1.0000 16.250 1.2297 0.10809 0.10255 -0.0628 0.0577 1.0000