XFOIL Version 6.96 Calculated polar for: GOE 679 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.250 -0.7987 0.11539 0.11258 -0.0619 1.0000 0.0300 -19.000 -0.8198 0.10871 0.10579 -0.0653 1.0000 0.0300 -18.750 -0.8370 0.10281 0.09980 -0.0682 1.0000 0.0301 -18.500 -0.8516 0.09746 0.09436 -0.0707 1.0000 0.0302 -18.250 -0.8661 0.09231 0.08913 -0.0731 1.0000 0.0303 -18.000 -0.8791 0.08747 0.08421 -0.0752 1.0000 0.0304 -17.750 -0.8928 0.08262 0.07928 -0.0774 1.0000 0.0305 -17.500 -0.9049 0.07831 0.07490 -0.0791 1.0000 0.0305 -17.250 -0.9188 0.07388 0.07039 -0.0809 1.0000 0.0305 -17.000 -0.9344 0.06940 0.06584 -0.0826 1.0000 0.0306 -16.750 -0.9502 0.06502 0.06138 -0.0841 1.0000 0.0307 -16.500 -0.9563 0.06037 0.05665 -0.0877 0.9994 0.0307 -16.250 -0.9594 0.05556 0.05174 -0.0922 0.9982 0.0308 -16.000 -0.9594 0.05106 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3.250 0.6187 0.00876 0.00368 -0.0461 0.3841 0.8359 3.500 0.7990 0.00919 0.00424 -0.0791 0.3697 0.9703 3.750 0.8374 0.00937 0.00441 -0.0813 0.3648 0.9840 4.000 0.8724 0.00961 0.00460 -0.0828 0.3588 0.9920 4.500 0.9566 0.01006 0.00494 -0.0890 0.3465 1.0000 4.750 0.9717 0.01018 0.00504 -0.0865 0.3410 1.0000 5.000 0.9850 0.01037 0.00518 -0.0836 0.3347 1.0000 5.250 1.0026 0.01048 0.00528 -0.0816 0.3302 1.0000 5.500 1.0198 0.01063 0.00541 -0.0795 0.3247 1.0000 5.750 1.0356 0.01085 0.00557 -0.0771 0.3184 1.0000 6.000 1.0538 0.01101 0.00572 -0.0753 0.3134 1.0000 6.250 1.0715 0.01120 0.00589 -0.0733 0.3077 1.0000 6.500 1.0874 0.01147 0.00611 -0.0711 0.3006 1.0000 6.750 1.1063 0.01167 0.00629 -0.0695 0.2942 1.0000 7.000 1.1243 0.01192 0.00651 -0.0677 0.2879 1.0000 7.250 1.1410 0.01222 0.00677 -0.0658 0.2806 1.0000 7.500 1.1595 0.01248 0.00701 -0.0642 0.2744 1.0000 7.750 1.1755 0.01285 0.00732 -0.0622 0.2666 1.0000 8.000 1.1948 0.01312 0.00759 -0.0608 0.2612 1.0000 8.250 1.2118 0.01347 0.00791 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1.0000 13.250 1.5265 0.02473 0.01925 -0.0313 0.1855 1.0000 13.500 1.5404 0.02557 0.02012 -0.0303 0.1834 1.0000 13.750 1.5521 0.02658 0.02114 -0.0291 0.1811 1.0000 14.000 1.5626 0.02767 0.02226 -0.0279 0.1787 1.0000 14.250 1.5691 0.02907 0.02367 -0.0265 0.1759 1.0000 14.500 1.5836 0.02994 0.02459 -0.0257 0.1746 1.0000 14.750 1.5966 0.03094 0.02564 -0.0249 0.1731 1.0000 15.000 1.6091 0.03198 0.02672 -0.0241 0.1710 1.0000 15.250 1.6189 0.03325 0.02803 -0.0232 0.1690 1.0000 15.500 1.6270 0.03466 0.02946 -0.0223 0.1664 1.0000 15.750 1.6308 0.03645 0.03127 -0.0212 0.1635 1.0000 16.000 1.6405 0.03780 0.03266 -0.0204 0.1614 1.0000 16.250 1.6509 0.03911 0.03403 -0.0198 0.1595 1.0000 16.500 1.6597 0.04057 0.03553 -0.0192 0.1568 1.0000 16.750 1.6640 0.04245 0.03744 -0.0185 0.1540 1.0000 17.000 1.6627 0.04488 0.03988 -0.0176 0.1504 1.0000 17.250 1.6716 0.04641 0.04147 -0.0172 0.1485 1.0000 17.500 1.6775 0.04826 0.04338 -0.0167 0.1469 1.0000 17.750 1.6803 0.05043 0.04558 -0.0163 0.1432 1.0000 18.000 1.6795 0.05298 0.04817 -0.0159 0.1404 1.0000 18.250 1.6775 0.05572 0.05095 -0.0155 0.1376 1.0000 18.500 1.6808 0.05794 0.05324 -0.0153 0.1348 1.0000 18.750 1.6796 0.06066 0.05600 -0.0152 0.1314 1.0000 19.000 1.6689 0.06448 0.05985 -0.0151 0.1272 1.0000 19.250 1.6675 0.06731 0.06274 -0.0152 0.1233 1.0000