XFOIL Version 6.94 Calculated polar for: agora06 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 -0.0728 0.03496 0.02379 -0.0208 0.9988 0.2768 -2.750 -0.0510 0.03543 0.02443 -0.0203 0.9988 0.2959 -2.500 -0.0287 0.03595 0.02497 -0.0200 0.9988 0.3156 -2.250 -0.0065 0.03647 0.02557 -0.0196 0.9988 0.3382 -2.000 0.1777 0.03319 0.02310 -0.0447 0.9059 0.4440 -1.750 0.1989 0.03406 0.02390 -0.0413 0.8765 0.4455 -1.500 0.2141 0.03511 0.02494 -0.0365 0.8455 0.4477 -1.250 0.2235 0.03645 0.02629 -0.0299 0.8076 0.4503 -1.000 0.2272 0.03826 0.02809 -0.0214 0.7602 0.4533 -0.750 0.2262 0.04058 0.03036 -0.0115 0.7001 0.4567 -0.500 0.2278 0.04289 0.03259 -0.0024 0.6254 0.4610 -0.250 0.2401 0.04273 0.03252 0.0042 0.5508 0.4677 0.000 0.2860 0.03728 0.02700 0.0023 0.4062 0.4837 0.250 0.3081 0.03704 0.02646 0.0053 0.3775 0.5043 0.500 0.3316 0.03677 0.02638 0.0073 0.3553 0.5626 0.750 0.3542 0.03548 0.02572 0.0103 0.3378 1.0012 1.000 0.3823 0.03628 0.02601 0.0115 0.3249 1.0012 1.250 0.4092 0.03723 0.02649 0.0128 0.3135 1.0012 1.500 0.4375 0.03830 0.02728 0.0135 0.3017 1.0012 1.750 0.4652 0.03948 0.02820 0.0144 0.2920 1.0012 2.000 0.4940 0.04079 0.02936 0.0150 0.2855 1.0012 2.250 0.5232 0.04227 0.03074 0.0153 0.2787 1.0012 2.500 0.5526 0.04391 0.03229 0.0153 0.2713 1.0012 2.750 0.5817 0.04571 0.03400 0.0153 0.2650 1.0012 3.000 0.6144 0.04788 0.03637 0.0140 0.2620 1.0012 3.250 0.6462 0.05034 0.03897 0.0126 0.2595 1.0012 3.500 0.6770 0.05310 0.04186 0.0109 0.2562 1.0012 3.750 0.7101 0.05681 0.04599 0.0066 0.2533 1.0012 4.000 0.7415 0.06145 0.05104 0.0016 0.2545 1.0012