XFOIL Version 6.96 Calculated polar for: NACA 64-108 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.6101 0.08462 0.08122 -0.0095 1.0000 0.0399 -8.500 -0.6168 0.07896 0.07562 -0.0147 1.0000 0.0402 -8.250 -0.6282 0.07375 0.07040 -0.0193 1.0000 0.0401 -8.000 -0.6344 0.06874 0.06530 -0.0225 1.0000 0.0410 -7.750 -0.6370 0.06409 0.06042 -0.0253 1.0000 0.0429 -7.500 -0.6359 0.06216 0.05797 -0.0261 1.0000 0.0443 -7.250 -0.6373 0.05434 0.04997 -0.0272 1.0000 0.0454 -7.000 -0.6238 0.05007 0.04581 -0.0274 1.0000 0.0471 -6.750 -0.6093 0.04751 0.04322 -0.0274 1.0000 0.0513 -6.250 -0.5815 0.04052 0.03570 -0.0272 1.0000 0.0620 -5.750 -0.5462 0.03476 0.02942 -0.0264 1.0000 0.0753 -5.500 -0.5277 0.03240 0.02671 -0.0258 1.0000 0.0871 -5.250 -0.4958 0.02549 0.01875 -0.0225 1.0000 0.0414 -5.000 -0.4728 0.02285 0.01576 -0.0214 1.0000 0.0405 -4.750 -0.4493 0.02092 0.01353 -0.0203 1.0000 0.0412 -4.500 -0.4259 0.01889 0.01126 -0.0190 1.0000 0.0404 -4.250 -0.4031 0.01735 0.00956 -0.0177 1.0000 0.0409 -4.000 -0.3812 0.01615 0.00826 -0.0164 1.0000 0.0424 -3.750 -0.3607 0.01513 0.00715 -0.0150 1.0000 0.0455 -3.500 -0.3420 0.01397 0.00605 -0.0135 1.0000 0.0495 -3.250 -0.3226 0.01323 0.00529 -0.0121 1.0000 0.0538 -3.000 -0.3033 0.01245 0.00449 -0.0109 1.0000 0.0608 -2.750 -0.2831 0.01183 0.00392 -0.0099 1.0000 0.0801 -2.500 -0.2690 0.00899 0.00339 -0.0089 1.0000 0.5905 -2.250 -0.2455 0.00882 0.00360 -0.0078 0.9969 0.7112 -2.000 -0.2087 0.00885 0.00370 -0.0094 0.9887 0.7643 -1.750 -0.1759 0.00903 0.00403 -0.0094 0.9811 0.8233 -1.500 -0.1477 0.00923 0.00431 -0.0082 0.9720 0.8658 -1.250 -0.1121 0.00926 0.00428 -0.0095 0.9647 0.8841 -1.000 -0.0733 0.00924 0.00421 -0.0116 0.9581 0.8963 -0.750 -0.0373 0.00922 0.00415 -0.0131 0.9504 0.9086 -0.500 0.0001 0.00917 0.00408 -0.0149 0.9432 0.9203 -0.250 0.0343 0.00914 0.00401 -0.0160 0.9339 0.9325 0.000 0.0722 0.00909 0.00395 -0.0178 0.9268 0.9438 0.250 0.1092 0.00905 0.00391 -0.0196 0.9171 0.9546 0.500 0.1507 0.00901 0.00387 -0.0223 0.9082 0.9628 0.750 0.1917 0.00895 0.00383 -0.0250 0.9000 0.9714 1.000 0.2343 0.00891 0.00381 -0.0281 0.8892 0.9791 1.250 0.2765 0.00886 0.00383 -0.0312 0.8786 0.9870 1.500 0.3174 0.00880 0.00381 -0.0340 0.8668 0.9962 1.750 0.3429 0.00876 0.00380 -0.0337 0.8508 1.0000 2.000 0.3561 0.00867 0.00366 -0.0303 0.8258 1.0000 2.250 0.3679 0.00851 0.00336 -0.0262 0.7868 1.0000 2.500 0.3838 0.00845 0.00321 -0.0232 0.7433 1.0000 2.750 0.4048 0.00852 0.00325 -0.0216 0.7144 1.0000 3.000 0.4261 0.00862 0.00326 -0.0200 0.6687 1.0000 3.250 0.4445 0.00901 0.00321 -0.0176 0.5500 1.0000 3.500 0.4487 0.01222 0.00406 -0.0148 0.0841 1.0000 3.750 0.4717 0.01319 0.00499 -0.0141 0.0608 1.0000 4.000 0.4951 0.01412 0.00597 -0.0133 0.0527 1.0000 4.250 0.5169 0.01546 0.00723 -0.0125 0.0457 1.0000 4.500 0.5417 0.01640 0.00825 -0.0118 0.0427 1.0000 4.750 0.5667 0.01768 0.00958 -0.0111 0.0407 1.0000 5.000 0.5926 0.01922 0.01123 -0.0104 0.0395 1.0000 5.250 0.6191 0.02099 0.01319 -0.0098 0.0388 1.0000 5.500 0.6435 0.02287 0.01512 -0.0095 0.0353 1.0000 5.750 0.6694 0.02506 0.01762 -0.0087 0.0358 1.0000 6.000 0.6941 0.02992 0.02325 -0.0068 0.0436 1.0000 6.250 0.7222 0.03732 0.03184 -0.0032 0.0804 1.0000 6.500 0.7421 0.03999 0.03454 -0.0028 0.0745 1.0000 6.750 0.7560 0.04253 0.03747 -0.0020 0.0632 1.0000 7.000 0.7745 0.04726 0.04193 -0.0026 0.0598 1.0000 7.250 0.7800 0.04950 0.04504 -0.0007 0.0522 1.0000 7.500 0.7920 0.05270 0.04841 -0.0003 0.0495 1.0000 7.750 0.8052 0.05654 0.05220 -0.0003 0.0477 1.0000 8.000 0.8035 0.06588 0.06156 -0.0012 0.0461 1.0000 8.250 0.8029 0.06814 0.06421 -0.0003 0.0457 1.0000 8.500 0.7895 0.07076 0.06731 -0.0004 0.0434 1.0000 8.750 0.7776 0.07513 0.07182 -0.0013 0.0428 1.0000 9.000 0.7610 0.07975 0.07650 -0.0033 0.0427 1.0000 9.250 0.7463 0.08573 0.08250 -0.0080 0.0429 1.0000