XFOIL Version 6.94 Calculated polar for: agora02 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0841 0.05940 0.03938 -0.0063 0.9988 1.0012 -2.750 -0.0763 0.05852 0.03774 -0.0081 0.9988 1.0012 -2.500 -0.0586 0.05833 0.03669 -0.0100 0.9988 1.0012 -2.250 -0.0384 0.05850 0.03608 -0.0113 0.9988 1.0012 -2.000 -0.0179 0.05886 0.03578 -0.0119 0.9988 1.0012 -1.750 0.0022 0.05934 0.03571 -0.0123 0.9988 1.0012 -1.500 0.0220 0.05991 0.03582 -0.0125 0.9988 1.0012 -1.250 0.0415 0.06057 0.03611 -0.0125 0.9988 1.0012 -1.000 0.0607 0.06131 0.03656 -0.0126 0.9988 1.0012 -0.750 0.0795 0.06214 0.03717 -0.0126 0.9988 1.0012 -0.500 0.0981 0.06309 0.03796 -0.0127 0.9988 1.0012 -0.250 0.1161 0.06416 0.03894 -0.0129 0.9988 1.0012 0.000 0.1335 0.06539 0.04014 -0.0132 0.9988 1.0012 0.250 0.1499 0.06682 0.04160 -0.0138 0.9988 1.0012 0.500 0.1646 0.06854 0.04340 -0.0146 0.9988 1.0012 0.750 0.1769 0.07066 0.04561 -0.0158 0.9988 1.0012 1.000 0.1859 0.07328 0.04831 -0.0176 0.9988 1.0012 1.250 0.1921 0.07636 0.05137 -0.0197 0.9988 1.0012 1.500 0.1974 0.07963 0.05457 -0.0220 0.9988 1.0012 1.750 0.2034 0.08290 0.05772 -0.0243 0.9988 1.0012 2.000 0.2102 0.08610 0.06079 -0.0265 0.9988 1.0012 2.250 0.2179 0.08923 0.06378 -0.0285 0.9988 1.0012 2.500 0.2263 0.09230 0.06673 -0.0304 0.9988 1.0012 2.750 0.2353 0.09537 0.06966 -0.0323 0.9988 1.0012 3.000 0.2449 0.09842 0.07258 -0.0342 0.9988 1.0012 3.250 0.2549 0.10145 0.07548 -0.0360 0.9988 1.0012 3.500 0.2651 0.10447 0.07836 -0.0377 0.9988 1.0012 3.750 0.2757 0.10748 0.08124 -0.0394 0.9988 1.0012 4.000 0.2864 0.11049 0.08412 -0.0411 0.9988 1.0012 4.250 0.2974 0.11349 0.08699 -0.0428 0.9988 1.0012 4.500 0.3085 0.11648 0.08987 -0.0445 0.9988 1.0012 4.750 0.3198 0.11948 0.09275 -0.0461 0.9988 1.0012 5.000 0.3313 0.12249 0.09563 -0.0478 0.9988 1.0012