XFOIL Version 6.96 Calculated polar for: NACA 63 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.500 -0.9281 0.10255 0.09878 -0.0532 1.0000 0.0242 -19.250 -0.9411 0.09683 0.09294 -0.0562 1.0000 0.0244 -19.000 -0.9484 0.09213 0.08815 -0.0586 1.0000 0.0245 -18.750 -0.9513 0.08818 0.08414 -0.0607 1.0000 0.0248 -18.500 -0.9543 0.08429 0.08020 -0.0627 1.0000 0.0250 -18.250 -0.9577 0.08036 0.07619 -0.0647 1.0000 0.0253 -18.000 -0.9613 0.07648 0.07222 -0.0666 1.0000 0.0255 -17.750 -0.9646 0.07274 0.06838 -0.0683 1.0000 0.0257 -17.500 -0.9673 0.06919 0.06473 -0.0699 1.0000 0.0259 -17.250 -0.9696 0.06572 0.06116 -0.0714 1.0000 0.0262 -17.000 -0.9700 0.06261 0.05794 -0.0726 1.0000 0.0265 -16.750 -0.9696 0.05970 0.05493 -0.0735 1.0000 0.0268 -16.500 -0.9685 0.05691 0.05203 -0.0745 1.0000 0.0271 -16.250 -0.9662 0.05433 0.04934 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3.000 0.6860 0.00976 0.00387 -0.0777 0.5584 0.6626 3.250 0.7147 0.00983 0.00395 -0.0779 0.5533 0.6683 3.500 0.7429 0.00990 0.00403 -0.0781 0.5484 0.6734 3.750 0.7711 0.00999 0.00416 -0.0782 0.5435 0.6786 4.000 0.7993 0.01004 0.00427 -0.0783 0.5381 0.6843 4.250 0.8274 0.01012 0.00434 -0.0784 0.5323 0.6897 4.500 0.8548 0.01020 0.00447 -0.0784 0.5260 0.6950 4.750 0.8823 0.01025 0.00458 -0.0784 0.5191 0.7010 5.000 0.9100 0.01038 0.00468 -0.0784 0.5130 0.7067 5.250 0.9370 0.01042 0.00482 -0.0783 0.5063 0.7119 5.500 0.9635 0.01051 0.00492 -0.0781 0.4977 0.7177 5.750 0.9899 0.01060 0.00504 -0.0779 0.4870 0.7240 6.000 1.0152 0.01072 0.00517 -0.0775 0.4771 0.7297 6.250 1.0416 0.01080 0.00534 -0.0773 0.4681 0.7357 6.500 1.0673 0.01098 0.00550 -0.0771 0.4591 0.7420 6.750 1.0923 0.01109 0.00568 -0.0766 0.4483 0.7478 7.000 1.1167 0.01126 0.00588 -0.0761 0.4358 0.7543 7.500 1.1620 0.01172 0.00635 -0.0746 0.4040 0.7673 7.750 1.1827 0.01204 0.00664 -0.0735 0.3840 0.7742 8.000 1.2003 0.01247 0.00700 -0.0720 0.3594 0.7812 8.250 1.2138 0.01302 0.00746 -0.0698 0.3297 0.7886 8.500 1.2239 0.01372 0.00801 -0.0672 0.2984 0.7968 8.750 1.2274 0.01440 0.00861 -0.0633 0.2701 0.8049 9.000 1.2291 0.01528 0.00937 -0.0595 0.2426 0.8146 9.250 1.2291 0.01624 0.01026 -0.0556 0.2169 0.8247 9.500 1.2300 0.01729 0.01124 -0.0522 0.1925 0.8362 9.750 1.2298 0.01850 0.01237 -0.0490 0.1709 0.8495 10.000 1.2305 0.01971 0.01356 -0.0460 0.1517 0.8651 10.250 1.2302 0.02098 0.01484 -0.0430 0.1342 0.8851 10.500 1.2281 0.02226 0.01614 -0.0398 0.1190 0.9196 10.750 1.2339 0.02380 0.01766 -0.0388 0.1032 1.0000 11.000 1.2386 0.02548 0.01927 -0.0376 0.0911 1.0000 11.250 1.2439 0.02715 0.02090 -0.0365 0.0817 1.0000 11.500 1.2492 0.02885 0.02258 -0.0354 0.0743 1.0000 11.750 1.2522 0.03079 0.02448 -0.0343 0.0682 1.0000 12.000 1.2597 0.03241 0.02613 -0.0336 0.0636 1.0000 12.250 1.2613 0.03456 0.02826 -0.0327 0.0592 1.0000 12.500 1.2693 0.03623 0.02998 -0.0321 0.0559 1.0000 12.750 1.2723 0.03839 0.03214 -0.0314 0.0529 1.0000 13.000 1.2769 0.04047 0.03426 -0.0309 0.0502 1.0000 13.250 1.2831 0.04247 0.03631 -0.0305 0.0479 1.0000 13.500 1.2854 0.04490 0.03875 -0.0302 0.0459 1.0000 13.750 1.2882 0.04734 0.04123 -0.0299 0.0440 1.0000 14.000 1.2947 0.04947 0.04343 -0.0298 0.0423 1.0000 14.250 1.2991 0.05186 0.04586 -0.0298 0.0407 1.0000 14.500 1.3000 0.05468 0.04870 -0.0298 0.0395 1.0000 14.750 1.3013 0.05753 0.05161 -0.0299 0.0383 1.0000 15.000 1.3079 0.05984 0.05400 -0.0301 0.0371 1.0000 15.250 1.3129 0.06238 0.05659 -0.0304 0.0359 1.0000 15.500 1.3161 0.06517 0.05942 -0.0308 0.0348 1.0000 15.750 1.3158 0.06840 0.06268 -0.0313 0.0339 1.0000 16.000 1.3174 0.07148 0.06582 -0.0318 0.0331 1.0000 16.250 1.3225 0.07419 0.06862 -0.0323 0.0323 1.0000 16.500 1.3265 0.07706 0.07157 -0.0330 0.0315 1.0000 16.750 1.3298 0.08006 0.07463 -0.0337 0.0307 1.0000 17.000 1.3325 0.08318 0.07780 -0.0345 0.0300 1.0000 17.250 1.3333 0.08658 0.08123 -0.0355 0.0294 1.0000