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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0827 0.03772 0.02625 -0.0187 0.9988 0.3980 -2.750 -0.0625 0.03796 0.02670 -0.0177 0.9988 0.4235 -2.500 -0.0422 0.03818 0.02712 -0.0167 0.9988 0.4530 -2.250 -0.0231 0.03836 0.02760 -0.0153 0.9988 0.4949 -2.000 -0.0038 0.03840 0.02800 -0.0138 0.9988 0.5326 -1.750 0.0221 0.03858 0.02815 -0.0141 0.9988 0.5419 -1.500 0.0491 0.03881 0.02833 -0.0148 0.9988 0.5476 -1.250 0.0759 0.03904 0.02863 -0.0154 0.9988 0.5551 -1.000 0.2749 0.03534 0.02576 -0.0386 0.8594 0.6146 -0.750 0.2764 0.03688 0.02774 -0.0286 0.8028 0.6496 -0.500 0.2685 0.03855 0.03009 -0.0161 0.7399 0.7406 -0.250 0.2722 0.04130 0.03283 -0.0067 0.6541 1.0012 0.000 0.3138 0.04248 0.03394 -0.0115 0.5134 1.0012 0.250 0.3450 0.04033 0.03125 -0.0088 0.4608 1.0012 0.500 0.3623 0.03884 0.02910 -0.0018 0.4402 1.0012 0.750 0.3812 0.03805 0.02763 0.0037 0.4199 1.0012 1.000 0.4031 0.03821 0.02714 0.0073 0.3990 1.0012 1.250 0.4286 0.03905 0.02751 0.0090 0.3803 1.0012 1.500 0.4554 0.04016 0.02827 0.0103 0.3675 1.0012 1.750 0.4841 0.04159 0.02948 0.0106 0.3552 1.0012 2.000 0.5109 0.04315 0.03078 0.0115 0.3435 1.0012 2.250 0.5411 0.04504 0.03266 0.0109 0.3336 1.0012 2.500 0.5712 0.04708 0.03466 0.0104 0.3278 1.0012 2.750 0.6026 0.04950 0.03715 0.0090 0.3213 1.0012 3.000 0.6345 0.05238 0.04019 0.0066 0.3143 1.0012 3.250 0.6680 0.05625 0.04439 0.0022 0.3099 1.0012 3.500 0.7017 0.06176 0.05035 -0.0048 0.3109 1.0012