XFOIL Version 6.96 Calculated polar for: NACA 64-208 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5528 0.08568 0.08343 -0.0165 1.0000 0.0150 -8.750 -0.5539 0.08123 0.07901 -0.0192 1.0000 0.0151 -8.500 -0.5573 0.07645 0.07426 -0.0227 1.0000 0.0154 -8.250 -0.5695 0.07018 0.06803 -0.0292 1.0000 0.0152 -8.000 -0.5747 0.06502 0.06281 -0.0333 1.0000 0.0153 -7.750 -0.5753 0.06020 0.05790 -0.0359 1.0000 0.0156 -7.500 -0.5722 0.05570 0.05328 -0.0376 1.0000 0.0159 -7.250 -0.5659 0.05145 0.04889 -0.0386 1.0000 0.0164 -7.000 -0.5568 0.04727 0.04453 -0.0390 1.0000 0.0173 -6.750 -0.5332 0.04534 0.04229 -0.0381 1.0000 0.0199 -6.500 -0.5210 0.04207 0.03874 -0.0374 1.0000 0.0200 -6.000 -0.5098 0.03012 0.02604 -0.0353 1.0000 0.0213 -5.750 -0.4956 0.02786 0.02367 -0.0341 1.0000 0.0222 -5.000 -0.4210 0.01836 0.01304 -0.0330 0.9958 0.0183 -4.750 -0.3863 0.01607 0.01043 -0.0344 0.9929 0.0187 -4.500 -0.3519 0.01465 0.00881 -0.0357 0.9892 0.0194 -4.250 -0.3175 0.01295 0.00695 -0.0373 0.9858 0.0212 -4.000 -0.2821 0.01156 0.00551 -0.0391 0.9830 0.0230 -3.750 -0.2481 0.01081 0.00469 -0.0406 0.9781 0.0248 -3.500 -0.2132 0.01016 0.00400 -0.0422 0.9730 0.0273 -3.250 -0.1770 0.00981 0.00362 -0.0441 0.9687 0.0304 -3.000 -0.1463 0.00899 0.00273 -0.0448 0.9597 0.0377 -2.750 -0.1144 0.00855 0.00226 -0.0457 0.9517 0.0511 -2.500 -0.0874 0.00718 0.00179 -0.0465 0.9419 0.2923 -2.250 -0.0634 0.00611 0.00164 -0.0463 0.9304 0.5613 -2.000 -0.0375 0.00589 0.00161 -0.0458 0.9195 0.6336 -1.750 -0.0112 0.00580 0.00156 -0.0453 0.9088 0.6707 -1.500 0.0149 0.00574 0.00153 -0.0447 0.8982 0.7068 -1.250 0.0400 0.00568 0.00158 -0.0438 0.8870 0.7506 -1.000 0.0655 0.00566 0.00161 -0.0430 0.8760 0.7797 -0.750 0.0913 0.00565 0.00162 -0.0423 0.8651 0.8005 -0.500 0.1178 0.00565 0.00159 -0.0419 0.8545 0.8138 -0.250 0.1444 0.00566 0.00157 -0.0414 0.8440 0.8256 0.000 0.1710 0.00565 0.00157 -0.0411 0.8329 0.8378 0.250 0.1976 0.00566 0.00158 -0.0407 0.8221 0.8499 0.500 0.2240 0.00568 0.00160 -0.0402 0.8115 0.8622 0.750 0.2501 0.00570 0.00162 -0.0397 0.8012 0.8750 1.000 0.2760 0.00572 0.00165 -0.0391 0.7896 0.8880 1.250 0.3012 0.00573 0.00168 -0.0384 0.7767 0.9010 1.500 0.3257 0.00575 0.00170 -0.0374 0.7612 0.9147 1.750 0.3490 0.00576 0.00168 -0.0361 0.7363 0.9296 2.000 0.3713 0.00580 0.00162 -0.0346 0.6997 0.9459 2.250 0.3980 0.00584 0.00161 -0.0342 0.6701 0.9632 2.500 0.4314 0.00593 0.00167 -0.0353 0.6429 0.9803 2.750 0.4645 0.00613 0.00171 -0.0366 0.5932 1.0000 3.000 0.4890 0.00656 0.00182 -0.0362 0.5085 1.0000 3.250 0.5078 0.00796 0.00221 -0.0354 0.2725 1.0000 3.500 0.5272 0.00960 0.00287 -0.0348 0.0585 1.0000 3.750 0.5533 0.01013 0.00335 -0.0347 0.0387 1.0000 4.000 0.5793 0.01071 0.00396 -0.0344 0.0312 1.0000 4.250 0.6053 0.01122 0.00451 -0.0342 0.0272 1.0000 4.500 0.6288 0.01222 0.00556 -0.0335 0.0241 1.0000 4.750 0.6521 0.01330 0.00673 -0.0328 0.0226 1.0000 5.000 0.6781 0.01380 0.00730 -0.0325 0.0211 1.0000 5.250 0.7027 0.01473 0.00830 -0.0320 0.0201 1.0000 5.500 0.7275 0.01578 0.00943 -0.0314 0.0192 1.0000 5.750 0.7524 0.01697 0.01071 -0.0309 0.0182 1.0000 6.000 0.7774 0.01837 0.01222 -0.0304 0.0175 1.0000