XFOIL Version 6.96 Calculated polar for: NACA 63(3)-618 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9337 0.09845 0.09521 -0.0633 1.0000 0.0152 -19.500 -0.9489 0.09246 0.08912 -0.0665 1.0000 0.0153 -19.250 -0.9621 0.08687 0.08345 -0.0695 1.0000 0.0154 -19.000 -0.9770 0.08114 0.07764 -0.0725 1.0000 0.0155 -18.750 -0.9868 0.07624 0.07265 -0.0751 1.0000 0.0157 -18.500 -0.9958 0.07161 0.06794 -0.0774 1.0000 0.0159 -18.250 -1.0017 0.06751 0.06376 -0.0795 1.0000 0.0160 -18.000 -1.0061 0.06370 0.05987 -0.0813 1.0000 0.0161 -17.750 -1.0082 0.06023 0.05634 -0.0829 1.0000 0.0165 -17.500 -1.0111 0.05681 0.05284 -0.0844 1.0000 0.0166 -17.250 -1.0114 0.05376 0.04972 -0.0857 1.0000 0.0169 -17.000 -1.0112 0.05086 0.04675 -0.0868 1.0000 0.0171 -16.750 -1.0062 0.04859 0.04442 -0.0876 1.0000 0.0174 -16.500 -1.0034 0.04612 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2.500 0.8100 0.00766 0.00250 -0.1176 0.5353 0.6084 2.750 0.8383 0.00775 0.00258 -0.1178 0.5306 0.6138 3.000 0.8669 0.00784 0.00267 -0.1180 0.5262 0.6189 3.250 0.8960 0.00787 0.00275 -0.1183 0.5220 0.6243 3.500 0.9245 0.00794 0.00284 -0.1185 0.5170 0.6296 3.750 0.9521 0.00808 0.00295 -0.1186 0.5113 0.6350 4.000 0.9809 0.00813 0.00303 -0.1189 0.5062 0.6398 4.250 1.0090 0.00820 0.00314 -0.1190 0.4997 0.6455 4.500 1.0361 0.00834 0.00326 -0.1190 0.4933 0.6514 4.750 1.0647 0.00840 0.00336 -0.1192 0.4879 0.6570 5.000 1.0921 0.00850 0.00349 -0.1192 0.4812 0.6629 5.250 1.1193 0.00862 0.00363 -0.1192 0.4746 0.6688 5.500 1.1467 0.00873 0.00375 -0.1193 0.4667 0.6745 5.750 1.1730 0.00889 0.00390 -0.1191 0.4577 0.6803 6.000 1.1994 0.00902 0.00406 -0.1190 0.4477 0.6866 6.250 1.2252 0.00920 0.00422 -0.1188 0.4358 0.6930 6.500 1.2499 0.00942 0.00442 -0.1184 0.4224 0.6990 6.750 1.2741 0.00966 0.00464 -0.1179 0.4098 0.7053 7.000 1.2987 0.00988 0.00485 -0.1174 0.3968 0.7118 7.250 1.3218 0.01015 0.00510 -0.1168 0.3813 0.7182 7.500 1.3426 0.01052 0.00542 -0.1158 0.3619 0.7250 7.750 1.3613 0.01099 0.00579 -0.1144 0.3365 0.7315 8.000 1.3760 0.01158 0.00626 -0.1124 0.3087 0.7383 8.250 1.3868 0.01224 0.00679 -0.1097 0.2810 0.7458 8.500 1.3946 0.01286 0.00732 -0.1064 0.2559 0.7530 8.750 1.3999 0.01363 0.00798 -0.1028 0.2313 0.7609 9.250 1.4089 0.01536 0.00954 -0.0961 0.1868 0.7775 9.500 1.4135 0.01633 0.01043 -0.0931 0.1670 0.7865 9.750 1.4171 0.01740 0.01145 -0.0901 0.1486 0.7961 10.000 1.4205 0.01858 0.01258 -0.0873 0.1322 0.8059 10.250 1.4233 0.01987 0.01382 -0.0846 0.1167 0.8168 10.500 1.4258 0.02127 0.01518 -0.0822 0.1016 0.8286 10.750 1.4297 0.02265 0.01656 -0.0800 0.0899 0.8419 11.000 1.4332 0.02412 0.01803 -0.0779 0.0792 0.8573 11.250 1.4358 0.02565 0.01958 -0.0758 0.0697 0.8773 11.500 1.4364 0.02717 0.02117 -0.0734 0.0617 0.9126 12.000 1.4433 0.03047 0.02453 -0.0702 0.0491 1.0000 12.250 1.4494 0.03222 0.02628 -0.0692 0.0448 1.0000 12.500 1.4541 0.03414 0.02820 -0.0681 0.0411 1.0000 12.750 1.4621 0.03582 0.02991 -0.0674 0.0389 1.0000 13.000 1.4671 0.03781 0.03192 -0.0666 0.0367 1.0000 13.250 1.4738 0.03973 0.03387 -0.0660 0.0351 1.0000 13.500 1.4814 0.04159 0.03577 -0.0654 0.0337 1.0000 13.750 1.4866 0.04374 0.03795 -0.0649 0.0323 1.0000 14.000 1.4900 0.04613 0.04037 -0.0644 0.0309 1.0000 14.250 1.4984 0.04805 0.04235 -0.0642 0.0301 1.0000 14.500 1.5052 0.05019 0.04454 -0.0640 0.0293 1.0000 14.750 1.5110 0.05246 0.04685 -0.0638 0.0283 1.0000 15.000 1.5145 0.05504 0.04947 -0.0637 0.0275 1.0000 15.250 1.5160 0.05793 0.05240 -0.0636 0.0265 1.0000 15.500 1.5235 0.06015 0.05469 -0.0637 0.0259 1.0000 15.750 1.5294 0.06260 0.05720 -0.0638 0.0252 1.0000 16.000 1.5343 0.06519 0.05984 -0.0640 0.0245 1.0000 16.250 1.5369 0.06813 0.06283 -0.0643 0.0238 1.0000 16.500 1.5369 0.07146 0.06621 -0.0647 0.0230 1.0000 16.750 1.5404 0.07438 0.06920 -0.0651 0.0224 1.0000 17.000 1.5458 0.07708 0.07197 -0.0656 0.0218 1.0000 17.250 1.5495 0.08004 0.07498 -0.0661 0.0211 1.0000 17.500 1.5513 0.08331 0.07831 -0.0668 0.0204 1.0000 17.750 1.5504 0.08701 0.08206 -0.0676 0.0197 1.0000