XFOIL Version 6.94 Calculated polar for: agora05 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1226 0.04082 0.03011 -0.0192 0.9992 0.5123 -2.750 -0.1103 0.03816 0.02863 -0.0128 0.9992 0.6595 -2.500 -0.0656 0.03516 0.02606 -0.0160 0.9992 1.0008 -2.250 -0.0330 0.03564 0.02561 -0.0185 0.9992 1.0008 -2.000 -0.0070 0.03617 0.02565 -0.0190 0.9992 1.0008 -1.750 0.0177 0.03674 0.02588 -0.0191 0.9992 1.0008 -1.500 0.0419 0.03734 0.02624 -0.0192 0.9992 1.0008 -1.250 0.0657 0.03799 0.02673 -0.0192 0.9992 1.0008 -1.000 0.0894 0.03868 0.02735 -0.0192 0.9992 1.0008 -0.750 0.1129 0.03946 0.02814 -0.0193 0.9992 1.0008 -0.500 0.1358 0.04042 0.02922 -0.0197 0.9992 1.0008 -0.250 0.3168 0.04688 0.03576 -0.0294 0.6774 1.0008 0.000 0.3290 0.05171 0.04043 -0.0258 0.6006 1.0008 0.250 0.3535 0.05433 0.04295 -0.0257 0.5514 1.0008 0.500 0.3975 0.05555 0.04431 -0.0355 0.5044 1.0008 0.750 0.4265 0.05618 0.04482 -0.0350 0.4831 1.0008 1.000 0.4527 0.05619 0.04464 -0.0318 0.4658 1.0008 1.250 0.4782 0.05603 0.04429 -0.0285 0.4493 1.0008 1.500 0.5045 0.05667 0.04478 -0.0266 0.4361 1.0008 1.750 0.5322 0.05867 0.04670 -0.0269 0.4266 1.0008 2.000 0.5593 0.06117 0.04916 -0.0280 0.4168 1.0008 2.250 0.5853 0.06433 0.05228 -0.0299 0.4075 1.0008 2.500 0.6085 0.07028 0.05830 -0.0361 0.4034 1.0008 2.750 0.6230 0.07821 0.06633 -0.0448 0.4070 1.0008 3.000 0.6363 0.08413 0.07225 -0.0495 0.4095 1.0008 3.250 0.6344 0.09177 0.07992 -0.0564 0.4175 1.0008 3.500 0.6283 0.09842 0.08656 -0.0613 0.4270 1.0008 3.750 0.6180 0.10453 0.09264 -0.0649 0.4395 1.0008 4.000 0.5837 0.11100 0.09907 -0.0678 0.4640 1.0008 4.250 0.5665 0.11740 0.10543 -0.0711 0.4940 1.0008 4.500 0.5474 0.12662 0.11470 -0.0795 0.5795 1.0008 4.750 0.5440 0.12948 0.11751 -0.0805 0.5922 1.0008 5.000 0.5551 0.13369 0.12168 -0.0823 0.5993 1.0008