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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.8636 0.10821 0.10266 -0.0534 1.0000 0.0362 -19.500 -0.8696 0.10356 0.09787 -0.0559 1.0000 0.0370 -19.250 -0.8744 0.09918 0.09334 -0.0582 1.0000 0.0376 -19.000 -0.8767 0.09532 0.08936 -0.0600 1.0000 0.0383 -18.750 -0.8743 0.09229 0.08633 -0.0615 1.0000 0.0390 -18.500 -0.8734 0.08904 0.08303 -0.0630 1.0000 0.0397 -18.250 -0.8721 0.08586 0.07978 -0.0645 1.0000 0.0405 -18.000 -0.8715 0.08261 0.07644 -0.0659 1.0000 0.0414 -17.750 -0.8693 0.07966 0.07337 -0.0671 1.0000 0.0423 -17.500 -0.8670 0.07674 0.07035 -0.0682 1.0000 0.0431 -17.250 -0.8623 0.07432 0.06793 -0.0691 1.0000 0.0439 -17.000 -0.8578 0.07184 0.06541 -0.0700 1.0000 0.0447 -16.750 -0.8533 0.06936 0.06289 -0.0708 1.0000 0.0457 -16.500 -0.8480 0.06702 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0.6171 3.000 0.6502 0.01510 0.00705 -0.0701 0.5396 0.6219 3.250 0.6775 0.01519 0.00718 -0.0701 0.5352 0.6260 3.500 0.7049 0.01528 0.00730 -0.0701 0.5311 0.6298 3.750 0.7335 0.01538 0.00738 -0.0703 0.5276 0.6343 4.000 0.7595 0.01552 0.00762 -0.0703 0.5230 0.6395 4.250 0.7863 0.01566 0.00780 -0.0703 0.5184 0.6450 4.500 0.8126 0.01578 0.00799 -0.0702 0.5143 0.6491 4.750 0.8402 0.01589 0.00812 -0.0702 0.5107 0.6537 5.000 0.8667 0.01605 0.00833 -0.0702 0.5067 0.6588 5.250 0.8921 0.01623 0.00859 -0.0700 0.5019 0.6642 5.500 0.9173 0.01637 0.00882 -0.0697 0.4974 0.6687 5.750 0.9437 0.01651 0.00899 -0.0696 0.4933 0.6737 6.000 0.9698 0.01668 0.00920 -0.0695 0.4891 0.6795 6.250 0.9933 0.01689 0.00950 -0.0690 0.4835 0.6854 6.500 1.0165 0.01700 0.00969 -0.0684 0.4770 0.6898 6.750 1.0398 0.01714 0.00987 -0.0677 0.4703 0.6951 7.000 1.0607 0.01734 0.01016 -0.0668 0.4628 0.7013 7.250 1.0843 0.01751 0.01034 -0.0663 0.4569 0.7073 7.500 1.1042 0.01773 0.01068 -0.0652 0.4505 0.7124 7.750 1.1237 0.01796 0.01099 -0.0640 0.4435 0.7182 8.000 1.1446 0.01818 0.01120 -0.0631 0.4369 0.7249 8.250 1.1589 0.01847 0.01163 -0.0611 0.4288 0.7306 8.500 1.1731 0.01871 0.01189 -0.0590 0.4210 0.7369 8.750 1.1840 0.01911 0.01237 -0.0567 0.4124 0.7439 9.000 1.1957 0.01952 0.01282 -0.0545 0.4036 0.7501 9.250 1.2048 0.02006 0.01346 -0.0521 0.3940 0.7570 9.500 1.2133 0.02068 0.01409 -0.0498 0.3828 0.7650 9.750 1.2176 0.02145 0.01493 -0.0472 0.3690 0.7720 10.000 1.2199 0.02242 0.01593 -0.0445 0.3521 0.7801 10.250 1.2203 0.02360 0.01708 -0.0419 0.3341 0.7881 10.500 1.2180 0.02501 0.01846 -0.0393 0.3156 0.7967 10.750 1.2155 0.02664 0.02004 -0.0370 0.2976 0.8062 11.000 1.2125 0.02837 0.02176 -0.0348 0.2809 0.8155 11.250 1.2101 0.03023 0.02360 -0.0329 0.2647 0.8260 11.500 1.2071 0.03219 0.02556 -0.0311 0.2496 0.8372 11.750 1.2041 0.03423 0.02760 -0.0294 0.2345 0.8495 12.000 1.2007 0.03637 0.02974 -0.0278 0.2199 0.8643 12.250 1.1978 0.03849 0.03190 -0.0264 0.2066 0.8840 12.500 1.1947 0.04070 0.03415 -0.0250 0.1939 0.9178 13.000 1.1916 0.04567 0.03909 -0.0242 0.1711 1.0000 13.250 1.1933 0.04822 0.04162 -0.0241 0.1603 1.0000 13.500 1.1951 0.05080 0.04419 -0.0242 0.1506 1.0000 13.750 1.1956 0.05358 0.04694 -0.0242 0.1415 1.0000 14.000 1.1975 0.05628 0.04964 -0.0244 0.1325 1.0000 14.250 1.1999 0.05897 0.05234 -0.0247 0.1240 1.0000 14.500 1.2004 0.06194 0.05530 -0.0250 0.1164 1.0000 14.750 1.2032 0.06472 0.05809 -0.0254 0.1088 1.0000 15.000 1.2051 0.06762 0.06101 -0.0259 0.1021 1.0000 15.250 1.2054 0.07078 0.06416 -0.0264 0.0957 1.0000 15.500 1.2085 0.07366 0.06707 -0.0270 0.0896 1.0000 15.750 1.2091 0.07690 0.07033 -0.0277 0.0842 1.0000 16.000 1.2110 0.08002 0.07347 -0.0285 0.0789 1.0000 16.250 1.2122 0.08327 0.07675 -0.0293 0.0747 1.0000 16.500 1.2137 0.08653 0.08004 -0.0302 0.0706 1.0000 16.750 1.2147 0.08990 0.08345 -0.0312 0.0672 1.0000 17.000 1.2169 0.09315 0.08675 -0.0322 0.0639 1.0000