XFOIL Version 6.96 Calculated polar for: NACA 64(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 -0.9089 0.09708 0.09407 -0.0703 1.0000 0.0177 -19.500 -0.9248 0.09089 0.08778 -0.0737 1.0000 0.0178 -19.250 -0.9364 0.08543 0.08222 -0.0767 1.0000 0.0179 -19.000 -0.9464 0.08037 0.07708 -0.0795 1.0000 0.0180 -18.750 -0.9526 0.07597 0.07260 -0.0818 1.0000 0.0181 -18.500 -0.9577 0.07181 0.06837 -0.0840 1.0000 0.0183 -18.250 -0.9630 0.06776 0.06425 -0.0861 1.0000 0.0184 -18.000 -0.9634 0.06441 0.06084 -0.0878 1.0000 0.0186 -17.750 -0.9664 0.06090 0.05725 -0.0894 1.0000 0.0188 -17.500 -0.9678 0.05768 0.05396 -0.0907 1.0000 0.0189 -17.250 -0.9669 0.05481 0.05102 -0.0919 1.0000 0.0191 -17.000 -0.9661 0.05203 0.04817 -0.0930 1.0000 0.0193 -16.750 -0.9647 0.04942 0.04548 -0.0938 1.0000 0.0195 -16.500 -0.9610 0.04712 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2.750 0.8321 0.00784 0.00266 -0.1181 0.5444 0.6119 3.000 0.8610 0.00789 0.00273 -0.1185 0.5416 0.6166 3.250 0.8897 0.00796 0.00281 -0.1187 0.5387 0.6210 3.500 0.9180 0.00803 0.00291 -0.1190 0.5357 0.6261 3.750 0.9458 0.00816 0.00303 -0.1191 0.5312 0.6313 4.000 0.9743 0.00817 0.00310 -0.1193 0.5272 0.6367 4.250 1.0024 0.00822 0.00317 -0.1195 0.5221 0.6416 4.500 1.0294 0.00832 0.00328 -0.1195 0.5161 0.6475 4.750 1.0572 0.00837 0.00338 -0.1196 0.5101 0.6535 5.000 1.0849 0.00844 0.00347 -0.1197 0.5048 0.6590 5.250 1.1116 0.00856 0.00360 -0.1196 0.4994 0.6653 5.500 1.1390 0.00862 0.00372 -0.1197 0.4928 0.6719 5.750 1.1647 0.00876 0.00384 -0.1194 0.4832 0.6781 6.000 1.1915 0.00885 0.00398 -0.1194 0.4752 0.6849 6.250 1.2172 0.00899 0.00414 -0.1192 0.4672 0.6917 6.500 1.2433 0.00912 0.00429 -0.1190 0.4591 0.6980 6.750 1.2672 0.00931 0.00448 -0.1185 0.4470 0.7049 7.000 1.2908 0.00952 0.00469 -0.1179 0.4311 0.7120 7.250 1.3089 0.00994 0.00499 -0.1163 0.3953 0.7187 7.500 1.3140 0.01088 0.00564 -0.1126 0.3395 0.7260 7.750 1.3151 0.01177 0.00629 -0.1081 0.2947 0.7333 8.000 1.3126 0.01277 0.00708 -0.1031 0.2515 0.7410 8.250 1.3137 0.01371 0.00787 -0.0990 0.2182 0.7493 8.500 1.3157 0.01467 0.00871 -0.0952 0.1910 0.7572 8.750 1.3187 0.01566 0.00960 -0.0918 0.1663 0.7661 9.000 1.3217 0.01673 0.01059 -0.0886 0.1451 0.7747 9.250 1.3252 0.01785 0.01166 -0.0857 0.1259 0.7845 9.500 1.3291 0.01904 0.01280 -0.0830 0.1101 0.7942 9.750 1.3335 0.02028 0.01401 -0.0806 0.0964 0.8051 10.000 1.3378 0.02159 0.01530 -0.0783 0.0840 0.8163 10.250 1.3436 0.02288 0.01659 -0.0763 0.0741 0.8280 10.750 1.3544 0.02566 0.01939 -0.0725 0.0570 0.8555 11.000 1.3599 0.02706 0.02084 -0.0708 0.0505 0.8732 11.250 1.3642 0.02846 0.02231 -0.0688 0.0454 0.8996 11.500 1.3664 0.02980 0.02378 -0.0666 0.0411 1.0000 11.750 1.3735 0.03146 0.02543 -0.0656 0.0373 1.0000 12.000 1.3824 0.03299 0.02699 -0.0647 0.0346 1.0000 12.250 1.3886 0.03480 0.02880 -0.0638 0.0320 1.0000 12.500 1.3976 0.03641 0.03044 -0.0631 0.0302 1.0000 12.750 1.4050 0.03820 0.03224 -0.0623 0.0284 1.0000 13.000 1.4116 0.04012 0.03419 -0.0617 0.0268 1.0000 13.250 1.4205 0.04185 0.03596 -0.0611 0.0254 1.0000 13.500 1.4265 0.04389 0.03803 -0.0606 0.0241 1.0000 13.750 1.4327 0.04598 0.04015 -0.0601 0.0229 1.0000 14.000 1.4413 0.04785 0.04206 -0.0598 0.0218 1.0000 14.250 1.4476 0.05000 0.04425 -0.0595 0.0208 1.0000 14.500 1.4505 0.05255 0.04683 -0.0591 0.0198 1.0000 14.750 1.4591 0.05454 0.04888 -0.0590 0.0191 1.0000 15.000 1.4666 0.05667 0.05105 -0.0589 0.0184 1.0000 15.250 1.4719 0.05910 0.05352 -0.0588 0.0177 1.0000 15.500 1.4753 0.06176 0.05622 -0.0588 0.0170 1.0000 15.750 1.4778 0.06462 0.05914 -0.0588 0.0166 1.0000 16.000 1.4847 0.06696 0.06154 -0.0590 0.0162 1.0000 16.250 1.4900 0.06955 0.06419 -0.0592 0.0157 1.0000 16.500 1.4956 0.07213 0.06683 -0.0595 0.0153 1.0000 16.750 1.4992 0.07500 0.06975 -0.0598 0.0150 1.0000 17.000 1.5028 0.07788 0.07269 -0.0603 0.0147 1.0000 17.250 1.5027 0.08136 0.07622 -0.0608 0.0143 1.0000 17.500 1.4991 0.08537 0.08031 -0.0616 0.0139 1.0000