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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.000 -0.7217 0.06595 0.05891 -0.0207 1.0000 0.1614 -7.750 -0.7430 0.06037 0.05261 -0.0202 1.0000 0.1405 -7.500 -0.7358 0.05503 0.04699 -0.0191 1.0000 0.1302 -7.250 -0.7274 0.05100 0.04251 -0.0175 1.0000 0.1233 -7.000 -0.7176 0.04730 0.03839 -0.0156 1.0000 0.1210 -6.750 -0.7060 0.04398 0.03462 -0.0137 1.0000 0.1204 -6.500 -0.6914 0.04084 0.03102 -0.0117 1.0000 0.1197 -6.250 -0.6733 0.03778 0.02747 -0.0098 1.0000 0.1184 -6.000 -0.6521 0.03504 0.02431 -0.0082 1.0000 0.1185 -5.750 -0.6289 0.03264 0.02158 -0.0069 1.0000 0.1220 -5.500 -0.6050 0.03051 0.01915 -0.0056 1.0000 0.1293 -5.250 -0.5759 0.02839 0.01693 -0.0050 1.0000 0.1383 -5.000 -0.5456 0.02641 0.01497 -0.0044 1.0000 0.1527 -4.750 -0.5209 0.02458 0.01336 -0.0032 1.0000 0.1773 -4.500 -0.5043 0.02260 0.01173 -0.0009 1.0000 0.2171 -4.250 -0.5102 0.01916 0.01043 0.0046 1.0000 0.4210 -4.000 -0.2856 0.02827 0.01888 -0.0003 1.0000 0.9198 -3.750 -0.2241 0.02717 0.01735 -0.0074 1.0000 0.9455 -3.500 -0.1756 0.02605 0.01594 -0.0128 1.0000 0.9658 -3.250 -0.1240 0.02475 0.01435 -0.0192 1.0000 0.9830 -3.000 -0.0723 0.02339 0.01278 -0.0258 1.0000 0.9986 -2.750 -0.0543 0.02274 0.01206 -0.0259 1.0000 1.0000 -2.500 -0.0400 0.02218 0.01145 -0.0252 1.0000 1.0000 -2.250 -0.0263 0.02167 0.01092 -0.0243 1.0000 1.0000 -2.000 -0.0133 0.02121 0.01045 -0.0232 1.0000 1.0000 -1.750 -0.0013 0.02080 0.01004 -0.0220 1.0000 1.0000 -1.500 0.0093 0.02044 0.00970 -0.0204 1.0000 1.0000 -1.250 0.0177 0.02014 0.00943 -0.0185 1.0000 1.0000 -1.000 0.0234 0.01989 0.00923 -0.0162 1.0000 1.0000 -0.750 0.0254 0.01973 0.00912 -0.0133 1.0000 1.0000 -0.500 0.0218 0.01965 0.00909 -0.0096 1.0000 1.0000 -0.250 0.0126 0.01963 0.00910 -0.0050 1.0000 1.0000 0.000 0.0000 0.01963 0.00911 0.0000 1.0000 1.0000 0.250 -0.0126 0.01963 0.00910 0.0050 1.0000 1.0000 0.500 -0.0219 0.01965 0.00909 0.0096 1.0000 1.0000 0.750 -0.0254 0.01973 0.00912 0.0133 1.0000 1.0000 1.000 -0.0234 0.01989 0.00923 0.0162 1.0000 1.0000 1.250 -0.0177 0.02013 0.00943 0.0185 1.0000 1.0000 1.500 -0.0092 0.02044 0.00970 0.0204 1.0000 1.0000 1.750 0.0013 0.02080 0.01004 0.0220 1.0000 1.0000 2.000 0.0133 0.02121 0.01044 0.0232 1.0000 1.0000 2.250 0.0263 0.02167 0.01091 0.0243 1.0000 1.0000 2.500 0.0401 0.02218 0.01144 0.0251 1.0000 1.0000 2.750 0.0544 0.02273 0.01205 0.0259 1.0000 1.0000 3.000 0.0723 0.02338 0.01277 0.0258 0.9987 1.0000 3.250 0.1239 0.02474 0.01434 0.0192 0.9832 1.0000 3.500 0.1755 0.02604 0.01592 0.0129 0.9659 1.0000 3.750 0.2239 0.02716 0.01733 0.0074 0.9458 1.0000 4.000 0.2857 0.02826 0.01887 0.0002 0.9199 1.0000 4.250 0.5098 0.01918 0.01042 -0.0045 0.4155 1.0000 4.500 0.5043 0.02260 0.01174 0.0009 0.2170 1.0000 4.750 0.5209 0.02459 0.01337 0.0032 0.1770 1.0000 5.000 0.5456 0.02641 0.01498 0.0044 0.1524 1.0000 5.250 0.5761 0.02840 0.01695 0.0050 0.1382 1.0000 5.500 0.6050 0.03051 0.01914 0.0056 0.1293 1.0000 5.750 0.6288 0.03264 0.02159 0.0069 0.1217 1.0000 6.000 0.6521 0.03504 0.02432 0.0082 0.1185 1.0000 6.250 0.6733 0.03779 0.02749 0.0098 0.1183 1.0000 6.500 0.6914 0.04084 0.03102 0.0117 0.1196 1.0000 6.750 0.7060 0.04398 0.03462 0.0137 0.1203 1.0000 7.000 0.7176 0.04730 0.03839 0.0157 0.1208 1.0000 7.250 0.7273 0.05100 0.04252 0.0175 0.1232 1.0000 7.500 0.7358 0.05505 0.04701 0.0191 0.1301 1.0000 7.750 0.7428 0.06033 0.05258 0.0202 0.1404 1.0000 8.000 0.7218 0.06595 0.05894 0.0207 0.1614 1.0000 8.250 0.7027 0.07421 0.06750 0.0176 0.2003 1.0000 8.500 0.5470 0.09473 0.08762 -0.0135 0.4031 1.0000 8.750 0.5513 0.09888 0.09174 -0.0142 0.4017 1.0000