XFOIL Version 6.94 Calculated polar for: Agora0 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1167 0.03849 0.02871 -0.0261 0.9992 0.3494 -2.750 -0.0910 0.03809 0.02827 -0.0262 0.9992 0.3612 -2.500 -0.0637 0.03780 0.02784 -0.0265 0.9992 0.3707 -2.250 0.1333 0.03322 0.02344 -0.0526 0.9065 0.4180 -2.000 0.1555 0.03388 0.02405 -0.0492 0.8779 0.4292 -1.750 0.1707 0.03454 0.02492 -0.0443 0.8480 0.4418 -1.500 0.1813 0.03557 0.02607 -0.0378 0.8108 0.4581 -1.250 0.1841 0.03709 0.02784 -0.0290 0.7662 0.4774 -1.000 0.1852 0.03869 0.02975 -0.0198 0.7073 0.5077 -0.750 0.1847 0.04002 0.03177 -0.0101 0.6396 0.5861 -0.250 0.2459 0.03404 0.02575 -0.0038 0.4132 1.0008 0.000 0.2676 0.03417 0.02502 -0.0001 0.3827 1.0008 0.250 0.2930 0.03473 0.02498 0.0018 0.3586 1.0008 0.500 0.3192 0.03541 0.02524 0.0033 0.3397 1.0008 0.750 0.3453 0.03614 0.02563 0.0049 0.3270 1.0008 1.000 0.3718 0.03702 0.02622 0.0062 0.3151 1.0008 1.250 0.3994 0.03805 0.02704 0.0072 0.3033 1.0008 1.500 0.4272 0.03918 0.02799 0.0081 0.2939 1.0008 1.750 0.4558 0.04044 0.02914 0.0089 0.2874 1.0008 2.000 0.4849 0.04185 0.03045 0.0094 0.2804 1.0008 2.250 0.5142 0.04341 0.03194 0.0096 0.2729 1.0008 2.500 0.5435 0.04514 0.03359 0.0098 0.2667 1.0008 2.750 0.5763 0.04720 0.03584 0.0087 0.2639 1.0008 3.000 0.6087 0.04954 0.03836 0.0074 0.2613 1.0008 3.250 0.6389 0.05204 0.04086 0.0065 0.2575 1.0008 3.500 0.6730 0.05550 0.04477 0.0026 0.2546 1.0008 3.750 0.7054 0.05969 0.04933 -0.0018 0.2553 1.0008