XFOIL Version 6.96 Calculated polar for: NACA 64(4)-221 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 -0.9404 0.10723 0.10142 -0.0425 1.0000 0.0318 -19.500 -0.9435 0.10308 0.09716 -0.0445 1.0000 0.0322 -19.250 -0.9451 0.09922 0.09317 -0.0464 1.0000 0.0325 -19.000 -0.9465 0.09547 0.08929 -0.0481 1.0000 0.0329 -18.750 -0.9467 0.09198 0.08566 -0.0496 1.0000 0.0333 -18.500 -0.9456 0.08870 0.08224 -0.0509 1.0000 0.0337 -18.250 -0.9433 0.08565 0.07907 -0.0521 1.0000 0.0341 -18.000 -0.9372 0.08329 0.07670 -0.0529 1.0000 0.0346 -17.750 -0.9318 0.08080 0.07419 -0.0538 1.0000 0.0351 -17.500 -0.9269 0.07825 0.07159 -0.0547 1.0000 0.0356 -17.250 -0.9212 0.07585 0.06913 -0.0554 1.0000 0.0361 -17.000 -0.9155 0.07343 0.06665 -0.0562 1.0000 0.0368 -16.750 -0.9094 0.07110 0.06423 -0.0568 1.0000 0.0376 -16.500 -0.9023 0.06895 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0.5630 0.6145 2.750 0.4647 0.01509 0.00694 -0.0368 0.5606 0.6187 3.000 0.4931 0.01517 0.00702 -0.0371 0.5582 0.6226 3.250 0.5216 0.01524 0.00714 -0.0373 0.5559 0.6257 3.500 0.5496 0.01535 0.00730 -0.0375 0.5534 0.6293 3.750 0.5760 0.01548 0.00754 -0.0375 0.5504 0.6334 4.000 0.6029 0.01562 0.00775 -0.0376 0.5476 0.6379 4.250 0.6302 0.01575 0.00793 -0.0377 0.5447 0.6422 4.500 0.6574 0.01586 0.00813 -0.0378 0.5421 0.6459 4.750 0.6851 0.01596 0.00831 -0.0379 0.5396 0.6502 5.000 0.7136 0.01607 0.00845 -0.0382 0.5373 0.6551 5.250 0.7410 0.01623 0.00865 -0.0383 0.5344 0.6603 5.500 0.7646 0.01637 0.00895 -0.0378 0.5295 0.6647 5.750 0.7900 0.01638 0.00905 -0.0375 0.5235 0.6694 6.000 0.8190 0.01632 0.00895 -0.0376 0.5177 0.6750 6.250 0.8399 0.01642 0.00919 -0.0367 0.5094 0.6812 6.500 0.8650 0.01635 0.00915 -0.0362 0.5012 0.6862 6.750 0.8866 0.01645 0.00939 -0.0353 0.4931 0.6919 7.000 0.9098 0.01653 0.00953 -0.0346 0.4857 0.6983 7.250 0.9330 0.01662 0.00968 -0.0340 0.4784 0.7041 7.500 0.9525 0.01675 0.00995 -0.0327 0.4691 0.7095 7.750 0.9719 0.01688 0.01015 -0.0315 0.4589 0.7159 8.000 0.9889 0.01702 0.01031 -0.0299 0.4453 0.7227 8.250 1.0001 0.01724 0.01064 -0.0273 0.4277 0.7284 8.500 1.0055 0.01753 0.01097 -0.0239 0.4042 0.7353 8.750 1.0050 0.01810 0.01140 -0.0198 0.3738 0.7431 9.000 0.9980 0.01909 0.01227 -0.0154 0.3446 0.7499 9.250 0.9912 0.02039 0.01345 -0.0116 0.3157 0.7583 9.500 0.9833 0.02194 0.01489 -0.0081 0.2881 0.7665 9.750 0.9740 0.02373 0.01659 -0.0048 0.2621 0.7751 10.000 0.9679 0.02556 0.01834 -0.0022 0.2382 0.7842 10.250 0.9616 0.02749 0.02022 0.0002 0.2166 0.7930 10.500 0.9572 0.02948 0.02215 0.0021 0.1972 0.8028 11.000 0.9527 0.03336 0.02598 0.0054 0.1642 0.8231 11.250 0.9524 0.03527 0.02788 0.0067 0.1499 0.8344 11.500 0.9538 0.03715 0.02977 0.0078 0.1369 0.8468 11.750 0.9558 0.03905 0.03167 0.0088 0.1248 0.8612 12.000 0.9589 0.04086 0.03352 0.0097 0.1144 0.8789 12.250 0.9631 0.04269 0.03540 0.0104 0.1055 0.9036 12.500 0.9728 0.04469 0.03745 0.0095 0.0963 0.9481 13.000 0.9842 0.04856 0.04130 0.0096 0.0825 1.0000 13.250 0.9908 0.05062 0.04335 0.0095 0.0766 1.0000 13.500 0.9967 0.05277 0.04548 0.0094 0.0715 1.0000 13.750 1.0043 0.05479 0.04752 0.0092 0.0667 1.0000 14.000 1.0099 0.05702 0.04974 0.0090 0.0628 1.0000 14.250 1.0172 0.05914 0.05189 0.0087 0.0590 1.0000 14.500 1.0228 0.06146 0.05420 0.0084 0.0560 1.0000 14.750 1.0301 0.06363 0.05643 0.0081 0.0533 1.0000 15.000 1.0350 0.06609 0.05889 0.0077 0.0508 1.0000 15.250 1.0416 0.06840 0.06124 0.0073 0.0486 1.0000 15.500 1.0477 0.07081 0.06370 0.0068 0.0466 1.0000 15.750 1.0518 0.07347 0.06638 0.0062 0.0449 1.0000 16.000 1.0579 0.07594 0.06890 0.0056 0.0434 1.0000 16.250 1.0636 0.07847 0.07149 0.0050 0.0418 1.0000 16.500 1.0683 0.08116 0.07422 0.0042 0.0405 1.0000 16.750 1.0714 0.08410 0.07718 0.0033 0.0394 1.0000 17.000 1.0772 0.08671 0.07988 0.0026 0.0382 1.0000 17.250 1.0822 0.08947 0.08271 0.0017 0.0371 1.0000 17.500 1.0861 0.09238 0.08567 0.0007 0.0361 1.0000 17.750 1.0896 0.09537 0.08870 -0.0003 0.0354 1.0000