XFOIL Version 6.96 Calculated polar for: NACA 4421 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9711 0.08521 0.08128 -0.0896 1.0000 0.0371 -19.500 -0.9981 0.07864 0.07460 -0.0926 1.0000 0.0373 -19.250 -1.0324 0.07136 0.06719 -0.0959 1.0000 0.0375 -19.000 -1.0643 0.06472 0.06044 -0.0986 1.0000 0.0376 -18.750 -1.0957 0.05846 0.05405 -0.1008 1.0000 0.0378 -18.500 -1.1250 0.05282 0.04830 -0.1024 1.0000 0.0380 -18.250 -1.1463 0.04782 0.04319 -0.1041 0.9996 0.0382 -18.000 -1.1511 0.04331 0.03857 -0.1075 0.9982 0.0385 -17.750 -1.1505 0.03948 0.03465 -0.1107 0.9965 0.0389 -17.500 -1.1451 0.03622 0.03131 -0.1136 0.9948 0.0394 -17.250 -1.1355 0.03350 0.02850 -0.1162 0.9934 0.0399 -17.000 -1.1223 0.03115 0.02608 -0.1187 0.9922 0.0404 -16.750 -1.1101 0.02918 0.02404 -0.1202 0.9892 0.0408 -16.500 -1.0963 0.02747 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0.8577 0.00998 0.00409 -0.0805 0.3916 0.6387 4.250 0.8815 0.01011 0.00427 -0.0798 0.3863 0.6667 4.500 0.9070 0.01019 0.00442 -0.0794 0.3822 0.6918 4.750 0.9318 0.01030 0.00459 -0.0789 0.3764 0.7141 5.000 0.9544 0.01049 0.00480 -0.0780 0.3684 0.7373 5.250 0.9781 0.01061 0.00500 -0.0773 0.3611 0.7653 5.500 0.9995 0.01081 0.00525 -0.0762 0.3515 0.7996 5.750 1.0207 0.01095 0.00551 -0.0750 0.3422 0.8450 6.000 1.0382 0.01119 0.00582 -0.0731 0.3310 0.8920 6.250 1.0555 0.01142 0.00611 -0.0711 0.3208 0.9362 6.500 1.0780 0.01174 0.00642 -0.0703 0.3107 0.9753 6.750 1.1052 0.01207 0.00670 -0.0707 0.3009 1.0000 7.000 1.1252 0.01241 0.00700 -0.0695 0.2931 1.0000 7.250 1.1442 0.01280 0.00734 -0.0683 0.2843 1.0000 7.500 1.1639 0.01318 0.00769 -0.0671 0.2765 1.0000 7.750 1.1830 0.01360 0.00807 -0.0660 0.2687 1.0000 8.000 1.2022 0.01403 0.00847 -0.0649 0.2625 1.0000 8.250 1.2222 0.01443 0.00886 -0.0639 0.2562 1.0000 8.500 1.2397 0.01495 0.00934 -0.0626 0.2493 1.0000 8.750 1.2591 0.01540 0.00978 -0.0616 0.2433 1.0000 9.000 1.2769 0.01593 0.01028 -0.0604 0.2365 1.0000 9.250 1.2948 0.01648 0.01081 -0.0593 0.2310 1.0000 9.500 1.3130 0.01702 0.01134 -0.0583 0.2251 1.0000 9.750 1.3286 0.01769 0.01198 -0.0570 0.2183 1.0000 10.000 1.3456 0.01831 0.01259 -0.0559 0.2119 1.0000 10.250 1.3590 0.01915 0.01338 -0.0544 0.2031 1.0000 10.500 1.3733 0.01995 0.01416 -0.0531 0.1944 1.0000 10.750 1.3861 0.02087 0.01504 -0.0516 0.1862 1.0000 11.000 1.3975 0.02187 0.01601 -0.0501 0.1777 1.0000 11.250 1.4100 0.02285 0.01697 -0.0488 0.1699 1.0000 11.500 1.4193 0.02407 0.01815 -0.0473 0.1612 1.0000 11.750 1.4272 0.02541 0.01945 -0.0457 0.1516 1.0000 12.000 1.4353 0.02679 0.02080 -0.0442 0.1430 1.0000 12.250 1.4439 0.02817 0.02216 -0.0429 0.1364 1.0000 12.500 1.4515 0.02967 0.02365 -0.0416 0.1302 1.0000 12.750 1.4600 0.03116 0.02514 -0.0405 0.1249 1.0000 13.000 1.4665 0.03283 0.02679 -0.0393 0.1196 1.0000 13.250 1.4747 0.03440 0.02838 -0.0383 0.1156 1.0000 13.500 1.4818 0.03610 0.03009 -0.0373 0.1119 1.0000 13.750 1.4889 0.03785 0.03185 -0.0364 0.1089 1.0000 14.000 1.4974 0.03952 0.03356 -0.0357 0.1066 1.0000 14.250 1.5046 0.04130 0.03537 -0.0350 0.1039 1.0000 14.500 1.5089 0.04340 0.03749 -0.0342 0.1014 1.0000 14.750 1.5154 0.04531 0.03943 -0.0336 0.0993 1.0000 15.000 1.5219 0.04728 0.04144 -0.0331 0.0974 1.0000 15.250 1.5267 0.04943 0.04363 -0.0325 0.0954 1.0000 15.500 1.5305 0.05169 0.04592 -0.0320 0.0934 1.0000 15.750 1.5328 0.05413 0.04840 -0.0316 0.0917 1.0000 16.000 1.5386 0.05624 0.05055 -0.0312 0.0902 1.0000 16.250 1.5415 0.05869 0.05305 -0.0309 0.0885 1.0000 16.500 1.5436 0.06123 0.05564 -0.0306 0.0867 1.0000 16.750 1.5440 0.06399 0.05843 -0.0304 0.0850 1.0000 17.000 1.5448 0.06671 0.06120 -0.0302 0.0835 1.0000 17.250 1.5476 0.06922 0.06377 -0.0301 0.0820 1.0000 17.500 1.5475 0.07211 0.06671 -0.0300 0.0803 1.0000 17.750 1.5474 0.07498 0.06961 -0.0300 0.0786 1.0000 18.000 1.5439 0.07832 0.07299 -0.0301 0.0769 1.0000 18.250 1.5459 0.08096 0.07569 -0.0301 0.0754 1.0000 18.500 1.5438 0.08416 0.07894 -0.0303 0.0736 1.0000 18.750 1.5421 0.08729 0.08211 -0.0305 0.0716 1.0000