XFOIL Version 6.96 Calculated polar for: GOE 514 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 ------ -------- --------- --------- -------- -------- -------- -18.000 -0.4346 0.12456 0.12236 -0.0843 0.9884 0.0125 -17.750 -0.4547 0.11590 0.11357 -0.0894 0.9876 0.0123 -17.500 -0.4703 0.10868 0.10627 -0.0941 0.9868 0.0125 -17.250 -0.4733 0.10357 0.10111 -0.0978 0.9861 0.0127 -17.000 -0.4783 0.09825 0.09574 -0.1016 0.9853 0.0128 -16.750 -0.4868 0.09232 0.08972 -0.1060 0.9847 0.0128 -16.500 -0.4886 0.08747 0.08482 -0.1099 0.9840 0.0128 -16.250 -0.4899 0.08269 0.08001 -0.1140 0.9835 0.0131 -16.000 -0.4953 0.07742 0.07466 -0.1184 0.9829 0.0130 -15.750 -0.5015 0.07304 0.07024 -0.1211 0.9807 0.0131 -15.500 -0.5129 0.06816 0.06530 -0.1240 0.9775 0.0131 -15.250 -0.5223 0.06270 0.05977 -0.1287 0.9754 0.0133 -15.000 -0.5360 0.05653 0.05352 -0.1346 0.9736 0.0132 -14.750 -0.5652 0.04759 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9.250 1.3383 0.01670 0.01130 -0.0575 0.2834 1.0000 9.500 1.3568 0.01694 0.01157 -0.0563 0.2807 1.0000 9.750 1.3731 0.01728 0.01193 -0.0547 0.2776 1.0000 10.000 1.3881 0.01769 0.01233 -0.0531 0.2739 1.0000 10.250 1.4008 0.01822 0.01284 -0.0511 0.2694 1.0000 10.500 1.4178 0.01857 0.01324 -0.0498 0.2671 1.0000 10.750 1.4349 0.01894 0.01362 -0.0485 0.2634 1.0000 11.000 1.4497 0.01942 0.01410 -0.0470 0.2592 1.0000 11.250 1.4626 0.02001 0.01469 -0.0453 0.2549 1.0000 11.500 1.4780 0.02049 0.01520 -0.0439 0.2516 1.0000 11.750 1.4941 0.02096 0.01570 -0.0427 0.2477 1.0000 12.000 1.5071 0.02159 0.01632 -0.0411 0.2424 1.0000 12.250 1.5189 0.02232 0.01706 -0.0395 0.2376 1.0000 12.500 1.5336 0.02291 0.01766 -0.0382 0.2315 1.0000 12.750 1.5425 0.02384 0.01857 -0.0364 0.2244 1.0000 13.000 1.5529 0.02472 0.01942 -0.0347 0.2143 1.0000 13.250 1.5603 0.02579 0.02047 -0.0329 0.2052 1.0000 13.500 1.5639 0.02715 0.02178 -0.0307 0.1920 1.0000 13.750 1.5628 0.02888 0.02342 -0.0282 0.1756 1.0000 14.000 1.5530 0.03129 0.02572 -0.0251 0.1555 1.0000 14.250 1.5470 0.03355 0.02792 -0.0227 0.1407 1.0000 14.500 1.5361 0.03633 0.03063 -0.0201 0.1253 1.0000 14.750 1.5268 0.03912 0.03338 -0.0179 0.1134 1.0000 15.000 1.5201 0.04180 0.03604 -0.0162 0.1052 1.0000 15.250 1.5114 0.04477 0.03900 -0.0146 0.0953 1.0000 15.500 1.4924 0.04884 0.04302 -0.0128 0.0787 1.0000 15.750 1.4529 0.05520 0.04930 -0.0108 0.0559 1.0000 16.000 1.4390 0.05924 0.05339 -0.0100 0.0492 1.0000 16.250 1.3957 0.06680 0.06097 -0.0092 0.0313 1.0000 16.500 1.3693 0.07279 0.06701 -0.0091 0.0218 1.0000 16.750 1.3499 0.07812 0.07242 -0.0093 0.0182 1.0000 17.000 1.3405 0.08225 0.07663 -0.0095 0.0175 1.0000 17.250 1.3255 0.08717 0.08164 -0.0100 0.0157 1.0000 17.500 1.3156 0.09150 0.08606 -0.0105 0.0152 1.0000