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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.8014 0.02966 0.02488 -0.0120 1.0000 0.0171 -7.750 -0.7918 0.02643 0.02126 -0.0091 1.0000 0.0169 -7.500 -0.7774 0.02405 0.01859 -0.0069 1.0000 0.0171 -7.250 -0.7610 0.02199 0.01627 -0.0049 1.0000 0.0173 -7.000 -0.7422 0.02080 0.01489 -0.0032 1.0000 0.0179 -6.750 -0.7223 0.02047 0.01444 -0.0017 1.0000 0.0186 -6.500 -0.7052 0.01804 0.01179 0.0002 1.0000 0.0192 -6.250 -0.6885 0.01635 0.01002 0.0022 1.0000 0.0200 -6.000 -0.6715 0.01545 0.00907 0.0041 1.0000 0.0206 -5.750 -0.6545 0.01483 0.00843 0.0060 1.0000 0.0217 -5.500 -0.6384 0.01422 0.00778 0.0081 1.0000 0.0225 -5.250 -0.6174 0.01368 0.00719 0.0092 0.9993 0.0240 -5.000 -0.5806 0.01322 0.00669 0.0071 0.9961 0.0257 -4.750 -0.5488 0.01200 0.00537 0.0058 0.9917 0.0282 -4.500 -0.5133 0.01143 0.00479 0.0038 0.9877 0.0315 -4.250 -0.4781 0.01100 0.00431 0.0020 0.9832 0.0361 -4.000 -0.4433 0.01045 0.00375 0.0003 0.9784 0.0462 -3.750 -0.4101 0.00968 0.00321 -0.0012 0.9734 0.0917 -3.500 -0.3821 0.00827 0.00266 -0.0024 0.9668 0.2857 -3.250 -0.3581 0.00704 0.00232 -0.0024 0.9576 0.5037 -3.000 -0.3252 0.00668 0.00218 -0.0035 0.9512 0.5758 -2.750 -0.2961 0.00650 0.00208 -0.0038 0.9401 0.6145 -2.500 -0.2663 0.00636 0.00199 -0.0041 0.9292 0.6432 -2.250 -0.2378 0.00626 0.00192 -0.0041 0.9176 0.6700 -2.000 -0.2106 0.00619 0.00187 -0.0037 0.9055 0.6956 -1.750 -0.1848 0.00614 0.00186 -0.0030 0.8929 0.7233 -1.500 -0.1590 0.00612 0.00184 -0.0024 0.8800 0.7426 -1.250 -0.1328 0.00610 0.00181 -0.0019 0.8675 0.7568 -1.000 -0.1064 0.00608 0.00177 -0.0014 0.8555 0.7689 -0.750 -0.0798 0.00607 0.00174 -0.0011 0.8438 0.7796 -0.500 -0.0532 0.00608 0.00171 -0.0007 0.8327 0.7903 -0.250 -0.0265 0.00607 0.00169 -0.0004 0.8219 0.8001 0.000 0.0000 0.00606 0.00169 0.0000 0.8107 0.8107 0.250 0.0266 0.00607 0.00169 0.0004 0.8001 0.8219 0.500 0.0532 0.00608 0.00171 0.0007 0.7903 0.8328 0.750 0.0799 0.00607 0.00174 0.0011 0.7796 0.8438 1.000 0.1065 0.00608 0.00177 0.0014 0.7689 0.8555 1.250 0.1328 0.00610 0.00181 0.0019 0.7568 0.8675 1.500 0.1590 0.00612 0.00184 0.0024 0.7426 0.8800 1.750 0.1848 0.00614 0.00186 0.0030 0.7232 0.8928 2.000 0.2106 0.00619 0.00187 0.0037 0.6957 0.9054 2.250 0.2378 0.00626 0.00192 0.0041 0.6698 0.9176 2.500 0.2663 0.00636 0.00199 0.0041 0.6430 0.9292 2.750 0.2961 0.00650 0.00208 0.0038 0.6146 0.9402 3.000 0.3253 0.00668 0.00218 0.0035 0.5769 0.9511 3.250 0.3581 0.00705 0.00232 0.0024 0.5034 0.9576 3.500 0.3823 0.00824 0.00266 0.0023 0.2911 0.9667 3.750 0.4101 0.00968 0.00322 0.0012 0.0917 0.9734 4.000 0.4433 0.01045 0.00375 -0.0003 0.0463 0.9784 4.250 0.4779 0.01104 0.00436 -0.0020 0.0355 0.9833 4.500 0.5132 0.01144 0.00480 -0.0038 0.0317 0.9877 4.750 0.5487 0.01202 0.00539 -0.0057 0.0280 0.9916 5.000 0.5808 0.01319 0.00666 -0.0071 0.0256 0.9961 5.250 0.6174 0.01367 0.00719 -0.0092 0.0239 0.9993 5.500 0.6386 0.01420 0.00775 -0.0082 0.0225 1.0000 5.750 0.6547 0.01480 0.00839 -0.0061 0.0215 1.0000 6.000 0.6715 0.01548 0.00910 -0.0041 0.0207 1.0000 6.250 0.6885 0.01637 0.01003 -0.0021 0.0199 1.0000 6.500 0.7053 0.01800 0.01175 -0.0002 0.0192 1.0000 6.750 0.7223 0.02048 0.01445 0.0017 0.0186 1.0000 7.000 0.7419 0.02102 0.01513 0.0033 0.0180 1.0000 7.250 0.7608 0.02209 0.01637 0.0049 0.0173 1.0000 7.500 0.7773 0.02413 0.01867 0.0069 0.0171 1.0000 7.750 0.7917 0.02646 0.02130 0.0091 0.0170 1.0000 8.000 0.8011 0.02971 0.02494 0.0120 0.0172 1.0000 8.250 0.8038 0.03390 0.02953 0.0154 0.0178 1.0000 14.500 0.5143 0.14813 0.14595 -0.0098 0.0186 1.0000 14.750 0.5160 0.15146 0.14928 -0.0119 0.0185 1.0000