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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1713 0.03740 0.02855 0.0262 0.9988 0.7772 -2.750 -0.1456 0.03711 0.02813 0.0238 0.9988 0.7851 -2.500 -0.1206 0.03693 0.02788 0.0216 0.9988 0.7972 -2.250 -0.0957 0.03681 0.02778 0.0199 0.9988 0.8158 -2.000 -0.0681 0.03676 0.02785 0.0175 0.9988 0.8467 -1.750 -0.0281 0.03689 0.02819 0.0120 0.9988 0.9168 -1.500 0.0044 0.03702 0.02811 0.0045 0.9988 1.0012 -1.250 0.0410 0.03811 0.02860 -0.0031 0.9988 1.0012 -1.000 0.0704 0.03964 0.02961 -0.0078 0.9988 1.0012 -0.500 0.3287 0.04945 0.03815 -0.0442 0.7486 1.0012 -0.250 0.3568 0.05429 0.04281 -0.0470 0.7190 1.0012 0.000 0.3840 0.05851 0.04691 -0.0513 0.6977 1.0012 0.250 0.4108 0.06269 0.05099 -0.0568 0.6920 1.0012 0.500 0.4308 0.06626 0.05454 -0.0640 0.7031 1.0012 0.750 0.4486 0.07030 0.05850 -0.0691 0.7169 1.0012 1.000 0.4311 0.07278 0.06093 -0.0725 0.7603 1.0012 1.250 0.3925 0.07483 0.06288 -0.0697 0.8102 1.0012 1.500 0.3549 0.07662 0.06458 -0.0658 0.8643 1.0012 1.750 0.3073 0.07752 0.06538 -0.0581 0.9221 1.0012 2.000 0.2816 0.07943 0.06717 -0.0528 0.9612 1.0012 2.250 0.2330 0.07939 0.06705 -0.0413 0.9988 1.0012 2.500 0.2455 0.08247 0.06999 -0.0429 0.9988 1.0012 2.750 0.2579 0.08558 0.07298 -0.0445 0.9988 1.0012 3.000 0.2703 0.08872 0.07600 -0.0461 0.9988 1.0012 3.250 0.2826 0.09189 0.07905 -0.0476 0.9988 1.0012 3.500 0.4138 0.10255 0.08962 -0.0715 0.8706 1.0012 3.750 0.5069 0.11227 0.09922 -0.0774 0.7135 1.0012 4.000 0.5459 0.11865 0.10549 -0.0759 0.6309 1.0012 4.250 0.5539 0.12231 0.10905 -0.0739 0.5940 1.0012 4.500 0.5731 0.12735 0.11401 -0.0737 0.5671 1.0012 4.750 0.5770 0.13076 0.11732 -0.0730 0.5506 1.0012 5.000 0.5851 0.13463 0.12111 -0.0729 0.5365 1.0012