XFOIL Version 6.96 Calculated polar for: WORTMANN FX 05-191 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 ------ -------- --------- --------- -------- -------- -------- -13.750 -0.2958 0.12635 0.12265 -0.0884 0.9759 0.0732 -13.500 -0.2871 0.12237 0.11866 -0.0913 0.9745 0.0757 -13.250 -0.3633 0.10345 0.09977 -0.1111 0.9723 0.0812 -13.000 -0.3176 0.10477 0.10112 -0.1056 0.9719 0.0830 -12.750 -0.3005 0.10383 0.10017 -0.1045 0.9694 0.0843 -12.500 -0.2087 0.08479 0.08125 -0.1215 0.9593 0.0964 -12.250 -0.1768 0.08446 0.08089 -0.1210 0.9587 0.0981 -11.500 -0.5047 0.03749 0.03219 -0.1275 0.9446 0.0486 -11.250 -0.5684 0.05371 0.04850 -0.1244 0.9435 0.0604 -11.000 -0.6118 0.04908 0.04334 -0.1143 0.9337 0.0480 -10.750 -0.6030 0.04551 0.03947 -0.1143 0.9312 0.0469 -10.500 -0.6254 0.04503 0.03887 -0.1069 0.9241 0.0466 -10.250 -0.6321 0.04332 0.03693 -0.1025 0.9188 0.0462 -10.000 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0.250 0.2644 0.02331 0.01714 -0.0647 0.7512 0.7592 0.500 0.3060 0.02294 0.01672 -0.0674 0.7502 0.7646 0.750 0.2907 0.02339 0.01721 -0.0602 0.7379 0.7717 1.000 0.3278 0.02312 0.01692 -0.0614 0.7366 0.7749 1.250 0.3307 0.02343 0.01723 -0.0580 0.7278 0.7837 1.500 0.3563 0.02325 0.01704 -0.0575 0.7241 0.7876 1.750 0.3897 0.02294 0.01673 -0.0580 0.7224 0.7908 2.000 0.4298 0.02255 0.01633 -0.0600 0.7212 0.7951 2.250 0.4800 0.02208 0.01581 -0.0648 0.7202 0.8019 2.500 0.5173 0.02170 0.01545 -0.0660 0.7190 0.8042 2.750 0.4917 0.02234 0.01615 -0.0566 0.7072 0.8108 3.000 0.5409 0.02188 0.01567 -0.0610 0.7059 0.8170 3.250 0.5815 0.02144 0.01524 -0.0630 0.7045 0.8199 3.500 0.5595 0.02207 0.01594 -0.0543 0.6939 0.8262 3.750 0.6029 0.02167 0.01554 -0.0574 0.6920 0.8312 4.000 0.6473 0.02120 0.01508 -0.0605 0.6903 0.8349 4.250 0.6889 0.02081 0.01471 -0.0626 0.6886 0.8378 4.500 0.6721 0.02141 0.01538 -0.0553 0.6786 0.8454 4.750 0.7188 0.02102 0.01499 -0.0591 0.6762 0.8495 5.000 0.7585 0.02060 0.01461 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1.1066 0.02131 0.01609 -0.0321 0.4678 0.9366 10.250 1.1072 0.02196 0.01670 -0.0286 0.4442 0.9407 10.500 1.1052 0.02297 0.01766 -0.0253 0.4127 0.9453 10.750 1.1000 0.02448 0.01898 -0.0221 0.3678 0.9493 11.000 1.0836 0.02649 0.02071 -0.0176 0.3169 0.9539 11.250 1.0651 0.02906 0.02297 -0.0136 0.2643 0.9590 11.500 1.0466 0.03205 0.02565 -0.0103 0.2136 0.9636 11.750 1.0323 0.03505 0.02839 -0.0078 0.1739 0.9684 12.000 1.0256 0.03782 0.03099 -0.0063 0.1457 0.9740 12.250 1.0250 0.04034 0.03341 -0.0057 0.1282 0.9815 12.500 1.0264 0.04263 0.03564 -0.0053 0.1160 1.0000 12.750 1.0341 0.04486 0.03787 -0.0055 0.1062 1.0000 13.000 1.0404 0.04722 0.04022 -0.0056 0.0984 1.0000 13.250 1.0481 0.04951 0.04250 -0.0059 0.0914 1.0000 13.500 1.0570 0.05165 0.04466 -0.0061 0.0851 1.0000 13.750 1.0670 0.05375 0.04676 -0.0065 0.0794 1.0000 14.000 1.0783 0.05564 0.04864 -0.0066 0.0742 1.0000 14.250 1.0914 0.05752 0.05058 -0.0071 0.0697 1.0000 14.500 1.1065 0.05892 0.05186 -0.0070 0.0650 1.0000 14.750 1.1194 0.06093 0.05405 -0.0078 0.0618 1.0000 15.000 1.1333 0.06276 0.05591 -0.0084 0.0585 1.0000 15.250 1.1531 0.06378 0.05682 -0.0084 0.0551 1.0000 15.500 1.1646 0.06599 0.05922 -0.0092 0.0529 1.0000 15.750 1.1775 0.06799 0.06133 -0.0099 0.0504 1.0000 16.000 1.1910 0.06988 0.06321 -0.0106 0.0483 1.0000 16.250 1.2089 0.07130 0.06466 -0.0108 0.0460 1.0000 16.500 1.2186 0.07375 0.06731 -0.0116 0.0443 1.0000 16.750 1.2290 0.07609 0.06978 -0.0123 0.0426 1.0000 17.000 1.2403 0.07830 0.07204 -0.0131 0.0411 1.0000 17.250 1.2605 0.07956 0.07324 -0.0134 0.0393 1.0000 17.500 1.2617 0.08303 0.07698 -0.0144 0.0382 1.0000 17.750 1.2650 0.08630 0.08046 -0.0154 0.0370 1.0000 18.000 1.2696 0.08945 0.08378 -0.0164 0.0359 1.0000 18.250 1.2746 0.09254 0.08700 -0.0176 0.0350 1.0000 18.500 1.2833 0.09511 0.08960 -0.0188 0.0341 1.0000 18.750 1.2943 0.09762 0.09216 -0.0193 0.0330 1.0000 19.000 1.2835 0.10283 0.09768 -0.0213 0.0326 1.0000 19.250 1.2717 0.10838 0.10353 -0.0237 0.0321 1.0000