XFOIL Version 6.96 Calculated polar for: GOE 5K 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.750 -0.5227 0.08978 0.08755 -0.0103 1.0000 0.0032 -7.500 -0.5248 0.08705 0.08484 -0.0106 1.0000 0.0032 -7.250 -0.5256 0.08394 0.08177 -0.0116 1.0000 0.0032 -7.000 -0.5221 0.08052 0.07837 -0.0137 1.0000 0.0033 -6.750 -0.5169 0.07690 0.07475 -0.0161 1.0000 0.0034 -6.500 -0.5098 0.07321 0.07106 -0.0185 1.0000 0.0034 -6.250 -0.5010 0.06938 0.06721 -0.0209 1.0000 0.0035 -6.000 -0.4905 0.06546 0.06326 -0.0231 1.0000 0.0035 -5.750 -0.4785 0.06157 0.05932 -0.0250 1.0000 0.0036 -5.500 -0.4649 0.05766 0.05535 -0.0267 1.0000 0.0037 -5.250 -0.4499 0.05378 0.05138 -0.0280 1.0000 0.0039 -5.000 -0.4243 0.04924 0.04671 -0.0312 0.9990 0.0041 -4.750 -0.3915 0.04452 0.04181 -0.0350 0.9971 0.0043 -4.500 -0.3557 0.03988 0.03691 -0.0380 0.9954 0.0045 -4.250 -0.3241 0.03583 0.03262 -0.0396 0.9932 0.0046 -3.750 -0.2674 0.02834 0.02465 -0.0430 0.9890 0.0050 -2.750 -0.1463 0.01676 0.01171 -0.0433 0.9806 0.0049 -2.500 -0.1150 0.01425 0.00885 -0.0431 0.9794 0.0028 -2.250 -0.0867 0.01231 0.00659 -0.0425 0.9774 0.0026 -2.000 -0.0594 0.01079 0.00479 -0.0417 0.9747 0.0030 -1.750 -0.0313 0.00977 0.00361 -0.0417 0.9724 0.0043 -1.500 -0.0020 0.00887 0.00256 -0.0419 0.9705 0.0048 -1.000 0.0556 0.00792 0.00151 -0.0423 0.9653 0.0338 -0.750 0.1017 0.00523 0.00167 -0.0475 0.9717 0.9915 -0.500 0.1362 0.00525 0.00158 -0.0492 0.9705 1.0000 0.000 0.1967 0.00521 0.00143 -0.0506 0.9638 1.0000 0.250 0.2283 0.00517 0.00138 -0.0516 0.9598 1.0000 0.500 0.2671 0.00508 0.00130 -0.0541 0.9558 1.0000 0.750 0.3036 0.00493 0.00117 -0.0560 0.9421 1.0000 1.000 0.3439 0.00479 0.00112 -0.0587 0.9239 1.0000 1.250 0.3906 0.00474 0.00098 -0.0625 0.8556 1.0000 1.500 0.3842 0.00692 0.00114 -0.0549 0.3705 1.0000 1.750 0.3939 0.00880 0.00168 -0.0518 0.0098 1.0000 2.000 0.4169 0.00925 0.00224 -0.0506 0.0054 1.0000 2.500 0.4601 0.01064 0.00383 -0.0477 0.0031 1.0000 2.750 0.4808 0.01172 0.00504 -0.0459 0.0031 1.0000 3.000 0.5028 0.01315 0.00663 -0.0442 0.0032 1.0000 7.000 0.7804 0.06646 0.06400 -0.0178 0.0028 1.0000 7.500 0.7141 0.06446 0.06240 -0.0158 0.0030 1.0000 8.000 0.7814 0.08422 0.08204 -0.0215 0.0027 1.0000 8.250 0.7728 0.08848 0.08631 -0.0237 0.0026 1.0000 8.500 0.7668 0.09320 0.09101 -0.0277 0.0026 1.0000