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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.5636 0.09582 0.08914 0.0195 1.0000 0.3869 -7.750 -0.5412 0.09158 0.08488 0.0214 1.0000 0.4179 -6.250 -0.5981 0.05266 0.04543 -0.0244 1.0000 0.1743 -6.000 -0.5780 0.04663 0.03834 -0.0267 1.0000 0.1368 -5.750 -0.5565 0.04252 0.03345 -0.0266 1.0000 0.1241 -5.500 -0.5350 0.03897 0.02947 -0.0261 1.0000 0.1231 -5.250 -0.5118 0.03571 0.02579 -0.0254 1.0000 0.1216 -5.000 -0.4866 0.03266 0.02230 -0.0246 1.0000 0.1184 -4.750 -0.4604 0.03004 0.01924 -0.0237 1.0000 0.1175 -4.500 -0.4337 0.02779 0.01666 -0.0226 1.0000 0.1193 -4.250 -0.4081 0.02584 0.01454 -0.0216 1.0000 0.1287 -4.000 -0.3817 0.02416 0.01274 -0.0202 1.0000 0.1390 -3.750 -0.3572 0.02252 0.01106 -0.0185 1.0000 0.1522 -3.500 -0.3366 0.02078 0.00962 -0.0170 1.0000 0.1839 -3.250 -0.1337 0.01944 0.00999 -0.0268 1.0000 1.0000 -3.000 -0.1179 0.01885 0.00922 -0.0265 1.0000 1.0000 -2.750 -0.1022 0.01833 0.00854 -0.0261 1.0000 1.0000 -2.500 -0.0871 0.01785 0.00794 -0.0255 1.0000 1.0000 -2.250 -0.0728 0.01743 0.00742 -0.0247 1.0000 1.0000 -2.000 -0.0596 0.01705 0.00693 -0.0237 1.0000 1.0000 -1.750 -0.0482 0.01673 0.00656 -0.0223 1.0000 1.0000 -1.500 -0.0398 0.01646 0.00626 -0.0203 1.0000 1.0000 -1.250 -0.0357 0.01624 0.00604 -0.0176 1.0000 1.0000 -1.000 -0.0368 0.01608 0.00588 -0.0140 1.0000 1.0000 -0.750 -0.0419 0.01596 0.00575 -0.0098 1.0000 1.0000 -0.500 -0.0462 0.01590 0.00564 -0.0056 1.0000 1.0000 -0.250 -0.0435 0.01591 0.00557 -0.0026 1.0000 1.0000 0.000 -0.0332 0.01601 0.00555 -0.0008 1.0000 1.0000 0.250 -0.0190 0.01617 0.00562 0.0003 1.0000 1.0000 0.500 -0.0027 0.01638 0.00576 0.0011 1.0000 1.0000 0.750 0.0146 0.01664 0.00597 0.0017 1.0000 1.0000 1.000 0.0327 0.01695 0.00623 0.0021 1.0000 1.0000 1.250 0.0510 0.01730 0.00656 0.0025 1.0000 1.0000 1.500 0.0696 0.01769 0.00696 0.0027 1.0000 1.0000 1.750 0.0882 0.01812 0.00740 0.0029 1.0000 1.0000 2.000 0.1069 0.01860 0.00791 0.0031 1.0000 1.0000 2.250 0.1256 0.01911 0.00848 0.0031 1.0000 1.0000 2.500 0.1442 0.01968 0.00911 0.0031 1.0000 1.0000 2.750 0.1627 0.02029 0.00983 0.0031 1.0000 1.0000 3.000 0.1810 0.02095 0.01060 0.0030 1.0000 1.0000 3.250 0.1991 0.02166 0.01144 0.0029 1.0000 1.0000 3.500 0.2452 0.02295 0.01302 -0.0027 0.9857 1.0000 3.750 0.3007 0.02436 0.01489 -0.0096 0.9648 1.0000 4.000 0.4933 0.01872 0.00891 -0.0096 0.3085 1.0000 4.250 0.5036 0.02210 0.01086 -0.0075 0.1814 1.0000 4.500 0.5265 0.02409 0.01262 -0.0061 0.1515 1.0000 4.750 0.5556 0.02598 0.01452 -0.0051 0.1346 1.0000 5.000 0.5851 0.02806 0.01661 -0.0045 0.1209 1.0000 5.250 0.6156 0.03043 0.01919 -0.0038 0.1163 1.0000 5.500 0.6444 0.03318 0.02229 -0.0031 0.1146 1.0000 5.750 0.6706 0.03660 0.02589 -0.0026 0.1123 1.0000 6.000 0.6936 0.03931 0.02925 -0.0015 0.1104 1.0000 6.250 0.7163 0.04309 0.03338 -0.0008 0.1118 1.0000 6.500 0.7351 0.04678 0.03813 0.0005 0.1213 1.0000 6.750 0.7520 0.05154 0.04359 0.0010 0.1343 1.0000 7.000 0.7645 0.05740 0.05021 0.0003 0.1588 1.0000 7.250 0.7612 0.06666 0.06030 -0.0058 0.2159 1.0000