XFOIL Version 6.96 Calculated polar for: FX 83-W-108 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 ------ -------- --------- --------- -------- -------- -------- -10.250 -0.3746 0.11248 0.10889 -0.0481 1.0000 0.0577 -10.000 -0.3874 0.11020 0.10669 -0.0477 1.0000 0.0579 -9.750 -0.4018 0.10852 0.10509 -0.0458 1.0000 0.0579 -9.500 -0.4190 0.10704 0.10369 -0.0433 1.0000 0.0580 -9.250 -0.4113 0.10302 0.09969 -0.0389 1.0000 0.0593 -9.000 -0.4023 0.10150 0.09817 -0.0368 0.9993 0.0606 -8.750 -0.3820 0.09834 0.09500 -0.0397 0.9967 0.0631 -8.500 -0.3682 0.09403 0.09069 -0.0449 0.9927 0.0662 -8.250 -0.3776 0.08572 0.08241 -0.0644 0.9811 0.0702 -8.000 -0.3699 0.07914 0.07585 -0.0682 0.9760 0.0715 -7.750 -0.3488 0.07710 0.07379 -0.0672 0.9726 0.0727 -7.500 -0.3252 0.07424 0.07091 -0.0697 0.9695 0.0748 -7.250 -0.3105 0.06996 0.06662 -0.0753 0.9616 0.0775 -7.000 -0.3254 0.03685 0.03157 -0.1131 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1.4045 0.04792 0.04077 -0.0520 0.0402 1.0000 12.750 1.4121 0.05018 0.04323 -0.0503 0.0394 1.0000 13.000 1.4174 0.05237 0.04559 -0.0486 0.0386 1.0000 13.250 1.4199 0.05475 0.04814 -0.0468 0.0379 1.0000 13.500 1.4239 0.05713 0.05061 -0.0454 0.0370 1.0000 13.750 1.4245 0.06085 0.05452 -0.0440 0.0364 1.0000 14.000 1.4150 0.06534 0.05931 -0.0422 0.0361 1.0000 14.250 1.3999 0.06993 0.06418 -0.0405 0.0359 1.0000 14.500 1.3834 0.07426 0.06877 -0.0391 0.0359 1.0000 14.750 1.3633 0.07911 0.07388 -0.0382 0.0359 1.0000 15.000 1.3417 0.08419 0.07920 -0.0379 0.0359 1.0000 15.250 1.3206 0.08927 0.08450 -0.0382 0.0359 1.0000 15.500 1.2963 0.09510 0.09056 -0.0393 0.0359 1.0000 15.750 1.2750 0.10071 0.09638 -0.0412 0.0361 1.0000 16.000 1.2498 0.10745 0.10332 -0.0441 0.0362 1.0000