XFOIL Version 6.94 Calculated polar for: agora10 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0689 0.03384 0.02378 -0.0134 0.9988 0.4188 -2.750 -0.0516 0.03429 0.02437 -0.0126 0.9988 0.4438 -2.500 0.0267 0.03323 0.02395 -0.0206 0.9579 0.5311 -2.250 0.0769 0.03371 0.02417 -0.0230 0.9044 0.5400 -2.000 0.1003 0.03472 0.02501 -0.0209 0.8685 0.5453 -1.750 0.1208 0.03586 0.02600 -0.0183 0.8369 0.5519 -1.500 0.1391 0.03694 0.02718 -0.0150 0.8057 0.5602 -1.250 0.1551 0.03852 0.02881 -0.0111 0.7714 0.5716 -1.000 0.1705 0.04054 0.03098 -0.0071 0.7225 0.5860 -0.750 0.1834 0.04456 0.03525 -0.0029 0.6528 0.6050 -0.500 0.2055 0.04972 0.04081 -0.0031 0.5755 0.6377 -0.250 0.2232 0.05275 0.04469 -0.0013 0.5276 0.7516 0.000 0.2561 0.05669 0.04840 -0.0042 0.4775 1.0012 0.250 0.2822 0.05987 0.05109 -0.0049 0.4340 1.0012 0.500 0.3121 0.06248 0.05335 -0.0066 0.4106 1.0012 0.750 0.3507 0.06435 0.05504 -0.0119 0.3928 1.0012 1.000 0.3874 0.06594 0.05648 -0.0166 0.3778 1.0012 1.250 0.4188 0.06848 0.05888 -0.0199 0.3692 1.0012 1.500 0.4481 0.07111 0.06139 -0.0229 0.3611 1.0012 1.750 0.4748 0.07340 0.06355 -0.0245 0.3524 1.0012 2.000 0.5010 0.07722 0.06734 -0.0291 0.3480 1.0012 2.250 0.5234 0.08209 0.07219 -0.0345 0.3496 1.0012 2.500 0.5426 0.08670 0.07677 -0.0387 0.3510 1.0012 2.750 0.5536 0.09242 0.08251 -0.0453 0.3570 1.0012 3.000 0.5615 0.09797 0.08805 -0.0501 0.3658 1.0012 3.750 0.5687 0.11564 0.10569 -0.0627 0.4195 1.0012 4.000 0.5813 0.12024 0.11023 -0.0637 0.4204 1.0012 4.250 0.5690 0.12209 0.11201 -0.0643 0.4228 1.0012 4.500 0.5708 0.12540 0.11525 -0.0653 0.4241 1.0012 4.750 0.5754 0.12889 0.11868 -0.0662 0.4247 1.0012 5.000 0.5808 0.13241 0.12214 -0.0672 0.4251 1.0012