XFOIL Version 6.96 Calculated polar for: GOE 383 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.250 -0.7552 0.11558 0.11284 -0.0549 1.0000 0.0254 -19.000 -0.8162 0.10111 0.09812 -0.0637 1.0000 0.0255 -18.750 -0.8510 0.09158 0.08843 -0.0697 0.9913 0.0255 -18.500 -0.8768 0.08368 0.08038 -0.0749 0.9784 0.0256 -18.250 -0.9007 0.07646 0.07301 -0.0794 0.9690 0.0256 -18.000 -0.9210 0.07014 0.06655 -0.0831 0.9613 0.0257 -17.750 -0.9392 0.06425 0.06052 -0.0865 0.9537 0.0257 -17.500 -0.9536 0.05889 0.05503 -0.0896 0.9473 0.0258 -17.250 -0.9655 0.05397 0.04998 -0.0925 0.9399 0.0258 -17.000 -0.9739 0.04954 0.04542 -0.0951 0.9331 0.0259 -16.750 -0.9796 0.04558 0.04134 -0.0974 0.9246 0.0259 -16.500 -0.9827 0.04210 0.03773 -0.0992 0.9160 0.0260 -16.250 -0.9835 0.03908 0.03458 -0.1006 0.9044 0.0260 -16.000 -0.9819 0.03643 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0.9206 0.00958 0.00413 -0.0787 0.3685 0.8344 4.250 0.9455 0.00966 0.00428 -0.0782 0.3650 0.8593 4.500 0.9692 0.00978 0.00445 -0.0774 0.3613 0.8866 4.750 0.9920 0.00992 0.00463 -0.0765 0.3578 0.9124 5.000 1.0166 0.01001 0.00479 -0.0758 0.3558 0.9387 5.250 1.0450 0.01015 0.00495 -0.0760 0.3529 0.9656 5.500 1.0791 0.01034 0.00514 -0.0776 0.3491 0.9830 5.750 1.1122 0.01057 0.00534 -0.0790 0.3448 0.9936 6.000 1.1423 0.01082 0.00556 -0.0799 0.3407 1.0000 6.250 1.1596 0.01095 0.00570 -0.0781 0.3384 1.0000 6.500 1.1794 0.01112 0.00587 -0.0767 0.3355 1.0000 6.750 1.1995 0.01134 0.00608 -0.0756 0.3313 1.0000 7.000 1.2180 0.01165 0.00635 -0.0742 0.3256 1.0000 7.250 1.2387 0.01192 0.00661 -0.0733 0.3204 1.0000 7.500 1.2594 0.01221 0.00689 -0.0724 0.3146 1.0000 7.750 1.2781 0.01261 0.00725 -0.0713 0.3086 1.0000 8.000 1.2987 0.01294 0.00757 -0.0704 0.3030 1.0000 8.250 1.3172 0.01337 0.00798 -0.0694 0.2956 1.0000 8.500 1.3355 0.01384 0.00842 -0.0683 0.2890 1.0000 8.750 1.3525 0.01439 0.00894 -0.0672 0.2803 1.0000 9.000 1.3680 0.01502 0.00953 -0.0659 0.2710 1.0000 9.250 1.3812 0.01581 0.01026 -0.0644 0.2607 1.0000 9.500 1.3939 0.01665 0.01106 -0.0629 0.2500 1.0000 9.750 1.4071 0.01751 0.01188 -0.0616 0.2411 1.0000 10.250 1.4323 0.01940 0.01373 -0.0591 0.2262 1.0000 10.750 1.4546 0.02160 0.01590 -0.0567 0.2120 1.0000 11.000 1.4621 0.02302 0.01729 -0.0554 0.2043 1.0000 11.250 1.4739 0.02418 0.01844 -0.0544 0.1987 1.0000 11.500 1.4828 0.02556 0.01981 -0.0534 0.1932 1.0000 11.750 1.4932 0.02688 0.02115 -0.0525 0.1894 1.0000 12.000 1.5051 0.02812 0.02241 -0.0518 0.1858 1.0000 12.250 1.5144 0.02958 0.02389 -0.0510 0.1819 1.0000 12.500 1.5212 0.03128 0.02559 -0.0501 0.1779 1.0000 12.750 1.5321 0.03266 0.02701 -0.0495 0.1746 1.0000 13.000 1.5384 0.03448 0.02883 -0.0487 0.1695 1.0000 13.250 1.5426 0.03649 0.03085 -0.0479 0.1647 1.0000 13.500 1.5491 0.03830 0.03267 -0.0472 0.1595 1.0000 13.750 1.5506 0.04062 0.03499 -0.0464 0.1543 1.0000 14.000 1.5559 0.04259 0.03699 -0.0457 0.1492 1.0000 14.250 1.5546 0.04522 0.03961 -0.0449 0.1428 1.0000 14.500 1.5536 0.04783 0.04222 -0.0442 0.1351 1.0000 14.750 1.5452 0.05118 0.04553 -0.0432 0.1234 1.0000 15.000 1.5148 0.05681 0.05105 -0.0417 0.0980 1.0000 15.250 1.4884 0.06223 0.05642 -0.0406 0.0808 1.0000 15.500 1.4780 0.06614 0.06033 -0.0401 0.0743 1.0000 15.750 1.4730 0.06954 0.06377 -0.0398 0.0709 1.0000 16.000 1.4697 0.07277 0.06704 -0.0397 0.0681 1.0000 16.250 1.4685 0.07584 0.07015 -0.0396 0.0661 1.0000 16.500 1.4665 0.07903 0.07340 -0.0396 0.0640 1.0000 16.750 1.4634 0.08237 0.07678 -0.0397 0.0621 1.0000 17.000 1.4627 0.08546 0.07992 -0.0398 0.0605 1.0000 17.250 1.4622 0.08853 0.08304 -0.0400 0.0587 1.0000 17.500 1.4577 0.09213 0.08669 -0.0403 0.0569 1.0000 17.750 1.4573 0.09524 0.08985 -0.0407 0.0553 1.0000 18.000 1.4542 0.09873 0.09339 -0.0411 0.0535 1.0000 18.250 1.4521 0.10205 0.09674 -0.0415 0.0515 1.0000 18.500 1.4482 0.10567 0.10042 -0.0421 0.0498 1.0000 18.750 1.4473 0.10889 0.10368 -0.0427 0.0478 1.0000