XFOIL Version 6.96 Calculated polar for: NACA 63(4)-421 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 ------ -------- --------- --------- -------- -------- -------- -19.000 -0.8146 0.10571 0.10097 -0.0606 1.0000 0.0514 -18.750 -0.8620 0.09466 0.08959 -0.0669 1.0000 0.0514 -18.500 -0.8886 0.08747 0.08216 -0.0708 1.0000 0.0518 -18.250 -0.9095 0.08147 0.07594 -0.0738 1.0000 0.0524 -18.000 -0.9306 0.07580 0.07002 -0.0765 1.0000 0.0531 -17.750 -0.9493 0.07077 0.06473 -0.0786 1.0000 0.0537 -17.500 -0.9680 0.06615 0.05983 -0.0805 1.0000 0.0542 -17.250 -0.9400 0.06586 0.05973 -0.0797 1.0000 0.0558 -17.000 -0.9366 0.06355 0.05737 -0.0803 1.0000 0.0570 -16.750 -0.9360 0.06093 0.05464 -0.0810 1.0000 0.0582 -16.500 -0.9399 0.05802 0.05156 -0.0817 1.0000 0.0594 -16.250 -0.9475 0.05497 0.04828 -0.0823 1.0000 0.0604 -16.000 -0.9268 0.05396 0.04736 -0.0819 1.0000 0.0621 -15.750 -0.9151 0.05256 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0.6854 3.500 0.6984 0.01922 0.01214 -0.0692 0.5804 0.6892 3.750 0.7213 0.01946 0.01250 -0.0686 0.5753 0.6935 4.000 0.7487 0.01955 0.01263 -0.0687 0.5712 0.6988 4.250 0.7804 0.01956 0.01258 -0.0696 0.5678 0.7050 4.500 0.8116 0.01961 0.01261 -0.0701 0.5648 0.7095 4.750 0.8299 0.01999 0.01318 -0.0688 0.5595 0.7143 5.000 0.8545 0.02019 0.01346 -0.0685 0.5547 0.7203 5.250 0.8848 0.02023 0.01349 -0.0692 0.5509 0.7269 5.500 0.9150 0.02020 0.01347 -0.0694 0.5476 0.7312 5.750 0.9380 0.02049 0.01387 -0.0689 0.5431 0.7365 6.000 0.9597 0.02076 0.01425 -0.0683 0.5375 0.7433 6.250 0.9877 0.02070 0.01421 -0.0683 0.5326 0.7491 6.500 1.0216 0.02045 0.01392 -0.0689 0.5280 0.7545 6.750 1.0379 0.02072 0.01435 -0.0674 0.5209 0.7613 7.000 1.0645 0.02067 0.01434 -0.0673 0.5150 0.7678 7.250 1.0955 0.02051 0.01417 -0.0675 0.5105 0.7733 7.500 1.1136 0.02078 0.01458 -0.0662 0.5042 0.7805 7.750 1.1362 0.02082 0.01469 -0.0655 0.4976 0.7876 8.000 1.1667 0.02060 0.01445 -0.0656 0.4921 0.7936 8.250 1.1817 0.02083 0.01483 -0.0639 0.4847 0.8014 8.500 1.2043 0.02073 0.01478 -0.0630 0.4774 0.8084 8.750 1.2270 0.02069 0.01478 -0.0620 0.4705 0.8159 9.000 1.2408 0.02083 0.01503 -0.0600 0.4620 0.8246 9.250 1.2634 0.02065 0.01484 -0.0589 0.4540 0.8318 9.500 1.2688 0.02085 0.01517 -0.0556 0.4438 0.8419 9.750 1.2778 0.02080 0.01516 -0.0524 0.4344 0.8508 10.000 1.2783 0.02105 0.01549 -0.0483 0.4242 0.8622 10.250 1.2813 0.02140 0.01590 -0.0448 0.4141 0.8744 10.500 1.2853 0.02171 0.01624 -0.0414 0.4037 0.8873 10.750 1.2805 0.02242 0.01709 -0.0374 0.3928 0.9042 11.000 1.2819 0.02296 0.01767 -0.0341 0.3819 0.9258 11.250 1.2835 0.02368 0.01849 -0.0319 0.3685 0.9901 11.500 1.2907 0.02510 0.01992 -0.0317 0.3522 1.0000 11.750 1.2953 0.02670 0.02146 -0.0312 0.3349 1.0000 12.000 1.2963 0.02854 0.02321 -0.0303 0.3170 1.0000 12.250 1.2948 0.03061 0.02520 -0.0292 0.2991 1.0000 12.500 1.2916 0.03287 0.02739 -0.0282 0.2814 1.0000 12.750 1.2873 0.03531 0.02975 -0.0271 0.2639 1.0000 13.000 1.2821 0.03794 0.03230 -0.0262 0.2469 1.0000 13.250 1.2767 0.04071 0.03501 -0.0255 0.2307 1.0000 13.500 1.2711 0.04364 0.03786 -0.0249 0.2153 1.0000 13.750 1.2653 0.04670 0.04085 -0.0244 0.2010 1.0000 14.000 1.2603 0.04981 0.04390 -0.0242 0.1876 1.0000 14.250 1.2568 0.05290 0.04695 -0.0241 0.1747 1.0000 14.500 1.2536 0.05604 0.05006 -0.0241 0.1627 1.0000 14.750 1.2500 0.05931 0.05329 -0.0243 0.1517 1.0000 15.000 1.2459 0.06270 0.05661 -0.0245 0.1415 1.0000 15.250 1.2426 0.06610 0.05998 -0.0249 0.1314 1.0000 15.500 1.2409 0.06944 0.06332 -0.0254 0.1218 1.0000 15.750 1.2389 0.07284 0.06669 -0.0259 0.1132 1.0000 16.250 1.2376 0.07948 0.07329 -0.0271 0.0991 1.0000 16.500 1.2381 0.08265 0.07640 -0.0277 0.0933 1.0000 16.750 1.2404 0.08572 0.07949 -0.0284 0.0882 1.0000 17.000 1.2438 0.08858 0.08231 -0.0291 0.0837 1.0000 17.250 1.2475 0.09152 0.08530 -0.0299 0.0797 1.0000 17.500 1.2532 0.09399 0.08768 -0.0304 0.0761 1.0000