XFOIL Version 6.96 Calculated polar for: NACA 64-008A AIRFOIL 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 ------ -------- --------- --------- -------- -------- -------- -9.250 -0.6837 0.08251 0.08034 -0.0050 1.0000 0.0157 -9.000 -0.6934 0.07649 0.07425 -0.0102 1.0000 0.0157 -8.750 -0.7046 0.07182 0.06948 -0.0125 1.0000 0.0157 -8.500 -0.7112 0.06726 0.06480 -0.0135 1.0000 0.0157 -8.000 -0.7390 0.05230 0.04940 -0.0148 1.0000 0.0164 -7.750 -0.7343 0.04821 0.04514 -0.0146 1.0000 0.0169 -7.500 -0.7238 0.04512 0.04192 -0.0142 1.0000 0.0173 -7.250 -0.7115 0.04213 0.03874 -0.0137 1.0000 0.0179 -7.000 -0.6975 0.03914 0.03554 -0.0129 1.0000 0.0187 -6.750 -0.6819 0.03615 0.03231 -0.0120 1.0000 0.0198 -6.000 -0.6309 0.02254 0.01708 -0.0054 1.0000 0.0163 -5.750 -0.6103 0.01969 0.01391 -0.0042 1.0000 0.0166 -5.500 -0.5884 0.01800 0.01211 -0.0034 1.0000 0.0179 -5.250 -0.5653 0.01616 0.01008 -0.0022 1.0000 0.0181 -5.000 -0.5423 0.01487 0.00866 -0.0012 1.0000 0.0188 -4.750 -0.5196 0.01386 0.00756 -0.0001 1.0000 0.0197 -4.500 -0.4973 0.01296 0.00660 0.0010 1.0000 0.0206 -4.250 -0.4748 0.01234 0.00594 0.0020 1.0000 0.0221 -4.000 -0.4526 0.01171 0.00526 0.0031 1.0000 0.0234 -3.750 -0.4304 0.01119 0.00468 0.0042 1.0000 0.0242 -3.500 -0.4104 0.01019 0.00357 0.0056 1.0000 0.0269 -3.250 -0.3881 0.00975 0.00310 0.0065 1.0000 0.0307 -3.000 -0.3649 0.00939 0.00271 0.0073 0.9999 0.0381 -2.750 -0.3311 0.00803 0.00213 0.0051 0.9961 0.2172 -2.500 -0.3007 0.00647 0.00180 0.0033 0.9917 0.5312 -2.250 -0.2660 0.00606 0.00171 0.0016 0.9861 0.6290 -2.000 -0.2300 0.00580 0.00170 -0.0002 0.9817 0.7003 -1.750 -0.1971 0.00566 0.00166 -0.0013 0.9736 0.7367 -1.500 -0.1650 0.00553 0.00167 -0.0020 0.9658 0.7831 -1.250 -0.1344 0.00544 0.00166 -0.0023 0.9561 0.8138 -1.000 -0.1058 0.00537 0.00163 -0.0022 0.9441 0.8329 -0.750 -0.0782 0.00532 0.00160 -0.0019 0.9314 0.8484 -0.500 -0.0516 0.00528 0.00156 -0.0014 0.9178 0.8621 -0.250 -0.0257 0.00525 0.00154 -0.0007 0.9038 0.8759 0.000 0.0000 0.00525 0.00153 0.0000 0.8899 0.8898 0.250 0.0257 0.00525 0.00154 0.0007 0.8759 0.9038 0.500 0.0516 0.00528 0.00156 0.0014 0.8621 0.9178 0.750 0.0782 0.00532 0.00160 0.0019 0.8484 0.9313 1.000 0.1058 0.00537 0.00163 0.0022 0.8329 0.9441 1.250 0.1344 0.00544 0.00166 0.0023 0.8138 0.9561 1.500 0.1650 0.00553 0.00167 0.0020 0.7830 0.9657 1.750 0.1971 0.00566 0.00166 0.0013 0.7368 0.9735 2.000 0.2300 0.00580 0.00170 0.0002 0.7002 0.9816 2.250 0.2660 0.00606 0.00171 -0.0016 0.6289 0.9861 2.500 0.3007 0.00646 0.00180 -0.0033 0.5316 0.9916 2.750 0.3310 0.00802 0.00213 -0.0051 0.2174 0.9961 3.000 0.3648 0.00939 0.00271 -0.0073 0.0381 0.9999 3.250 0.3880 0.00974 0.00310 -0.0065 0.0307 1.0000 3.500 0.4103 0.01019 0.00357 -0.0056 0.0269 1.0000 3.750 0.4302 0.01119 0.00468 -0.0041 0.0242 1.0000 4.000 0.4525 0.01171 0.00526 -0.0030 0.0234 1.0000 4.250 0.4746 0.01234 0.00594 -0.0019 0.0222 1.0000 4.500 0.4972 0.01297 0.00661 -0.0009 0.0206 1.0000 4.750 0.5195 0.01385 0.00756 0.0002 0.0197 1.0000 5.000 0.5422 0.01487 0.00866 0.0012 0.0188 1.0000 5.250 0.5652 0.01616 0.01008 0.0023 0.0181 1.0000 5.500 0.5884 0.01803 0.01214 0.0034 0.0180 1.0000 5.750 0.6101 0.01974 0.01397 0.0042 0.0165 1.0000 6.000 0.6309 0.02254 0.01708 0.0054 0.0163 1.0000 11.750 0.5441 0.12552 0.12335 -0.0100 0.0156 1.0000 12.000 0.5437 0.12950 0.12733 -0.0115 0.0156 1.0000