XFOIL Version 6.96 Calculated polar for: NACA 64A-010 10.0% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.5270 0.09029 0.08562 -0.0143 1.0000 0.1539 -9.500 -0.5136 0.08673 0.08205 -0.0127 1.0000 0.1622 -9.250 -0.5425 0.08101 0.07642 -0.0168 1.0000 0.1683 -9.000 -0.5878 0.07366 0.06916 -0.0231 1.0000 0.1691 -8.750 -0.6333 0.06888 0.06438 -0.0246 1.0000 0.1694 -8.500 -0.6868 0.07377 0.06901 -0.0206 1.0000 0.1652 -7.250 -0.7322 0.04505 0.03821 -0.0163 1.0000 0.0886 -7.000 -0.7197 0.04105 0.03323 -0.0124 1.0000 0.0713 -6.750 -0.7056 0.03752 0.02938 -0.0105 1.0000 0.0707 -6.500 -0.6890 0.03409 0.02568 -0.0087 1.0000 0.0687 -6.250 -0.6707 0.03120 0.02241 -0.0067 1.0000 0.0671 -6.000 -0.6502 0.02877 0.01962 -0.0049 1.0000 0.0665 -5.750 -0.6282 0.02672 0.01728 -0.0034 1.0000 0.0678 -5.500 -0.6065 0.02543 0.01567 -0.0018 1.0000 0.0713 -5.250 -0.5825 0.02319 0.01340 -0.0009 1.0000 0.0749 -5.000 -0.5601 0.02169 0.01193 0.0003 1.0000 0.0792 -4.750 -0.5390 0.02061 0.01075 0.0020 1.0000 0.0854 -4.500 -0.5228 0.01926 0.00960 0.0040 1.0000 0.0959 -4.250 -0.5089 0.01810 0.00855 0.0066 1.0000 0.1084 -4.000 -0.4973 0.01692 0.00756 0.0095 1.0000 0.1354 -3.750 -0.5037 0.01389 0.00646 0.0147 1.0000 0.4137 -3.500 -0.5047 0.01335 0.00695 0.0221 1.0000 0.6733 -3.250 -0.4933 0.01355 0.00726 0.0266 1.0000 0.7367 -3.000 -0.4805 0.01379 0.00751 0.0308 1.0000 0.7795 -2.750 -0.4670 0.01408 0.00780 0.0349 1.0000 0.8164 -2.500 -0.4521 0.01462 0.00837 0.0396 1.0000 0.8572 -2.250 -0.4096 0.01588 0.00950 0.0405 1.0000 0.9011 -2.000 -0.3598 0.01631 0.00974 0.0368 1.0000 0.9185 -1.750 -0.3100 0.01647 0.00970 0.0322 1.0000 0.9295 -1.500 -0.2634 0.01652 0.00961 0.0278 1.0000 0.9400 -1.250 -0.2213 0.01655 0.00953 0.0240 1.0000 0.9516 -1.000 -0.1710 0.01656 0.00943 0.0185 1.0000 0.9604 -0.750 -0.1250 0.01655 0.00936 0.0137 1.0000 0.9709 -0.500 -0.0782 0.01654 0.00931 0.0085 1.0000 0.9818 -0.250 -0.0294 0.01651 0.00925 0.0028 1.0000 0.9926 0.000 0.0000 0.01649 0.00922 0.0000 1.0000 1.0000 0.250 0.0293 0.01651 0.00925 -0.0028 0.9926 1.0000 0.500 0.0782 0.01654 0.00930 -0.0085 0.9818 1.0000 0.750 0.1251 0.01655 0.00936 -0.0137 0.9709 1.0000 1.000 0.1709 0.01656 0.00943 -0.0185 0.9605 1.0000 1.250 0.2215 0.01654 0.00953 -0.0240 0.9516 1.0000 1.500 0.2633 0.01652 0.00961 -0.0277 0.9400 1.0000 1.750 0.3101 0.01647 0.00972 -0.0322 0.9295 1.0000 2.000 0.3598 0.01631 0.00973 -0.0368 0.9185 1.0000 2.250 0.4092 0.01589 0.00951 -0.0404 0.9013 1.0000 2.500 0.4522 0.01464 0.00838 -0.0397 0.8578 1.0000 2.750 0.4673 0.01409 0.00781 -0.0350 0.8169 1.0000 3.000 0.4806 0.01379 0.00751 -0.0308 0.7798 1.0000 3.250 0.4934 0.01355 0.00726 -0.0267 0.7368 1.0000 3.500 0.5048 0.01335 0.00695 -0.0221 0.6735 1.0000 3.750 0.5039 0.01389 0.00646 -0.0148 0.4148 1.0000 4.000 0.4975 0.01691 0.00756 -0.0095 0.1358 1.0000 4.250 0.5090 0.01811 0.00856 -0.0066 0.1081 1.0000 4.500 0.5228 0.01926 0.00961 -0.0040 0.0960 1.0000 4.750 0.5391 0.02060 0.01075 -0.0020 0.0852 1.0000 5.000 0.5602 0.02170 0.01194 -0.0003 0.0791 1.0000 5.250 0.5826 0.02322 0.01343 0.0009 0.0747 1.0000 5.500 0.6067 0.02547 0.01570 0.0018 0.0714 1.0000 5.750 0.6282 0.02673 0.01729 0.0034 0.0675 1.0000 6.000 0.6503 0.02877 0.01963 0.0049 0.0666 1.0000 6.250 0.6707 0.03121 0.02243 0.0067 0.0670 1.0000 6.500 0.6891 0.03409 0.02568 0.0087 0.0687 1.0000 6.750 0.7054 0.03746 0.02934 0.0105 0.0705 1.0000 7.000 0.7198 0.04115 0.03333 0.0123 0.0714 1.0000 7.250 0.7290 0.04481 0.03790 0.0162 0.0826 1.0000 8.750 0.7267 0.07797 0.07308 0.0226 0.1550 1.0000 9.000 0.6732 0.08129 0.07648 0.0214 0.1534 1.0000 9.250 0.6354 0.08764 0.08275 0.0144 0.1497 1.0000 9.500 0.6791 0.08924 0.08441 0.0220 0.1406 1.0000 9.750 0.6330 0.09614 0.09117 0.0123 0.1390 1.0000 10.000 0.6221 0.10064 0.09561 0.0090 0.1320 1.0000