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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2497 0.05096 0.03916 0.0366 0.9988 0.7971 -2.750 -0.2127 0.04929 0.03768 0.0337 0.9988 0.8389 -2.250 -0.0413 0.04585 0.03461 0.0012 0.9988 1.0012 -2.000 -0.0584 0.04379 0.03242 0.0018 0.9988 1.0012 -1.750 -0.0253 0.04298 0.03064 -0.0078 0.9988 1.0012 -1.500 0.0132 0.04355 0.03013 -0.0131 0.9988 1.0012 -1.250 0.0416 0.04428 0.03022 -0.0147 0.9988 1.0012 -1.000 0.0662 0.04507 0.03063 -0.0153 0.9988 1.0012 -0.750 0.0890 0.04597 0.03131 -0.0157 0.9988 1.0012 -0.500 0.1103 0.04703 0.03229 -0.0161 0.9988 1.0012 -0.250 0.1295 0.04843 0.03375 -0.0170 0.9988 1.0012 0.000 0.1422 0.05087 0.03641 -0.0190 0.9988 1.0012 0.250 0.1393 0.05559 0.04128 -0.0234 0.9988 1.0012 0.500 0.2431 0.06441 0.05002 -0.0502 0.9263 1.0012 0.750 0.2933 0.07002 0.05551 -0.0601 0.8867 1.0012 1.000 0.3470 0.07569 0.06106 -0.0679 0.8341 1.0012 1.250 0.4022 0.08159 0.06684 -0.0732 0.7693 1.0012 1.500 0.4026 0.08510 0.07023 -0.0733 0.7688 1.0012 1.750 0.4011 0.08844 0.07345 -0.0738 0.7796 1.0012 2.000 0.3922 0.09120 0.07609 -0.0736 0.7998 1.0012 2.250 0.3791 0.09351 0.07829 -0.0728 0.8284 1.0012 2.500 0.3564 0.09495 0.07962 -0.0702 0.8669 1.0012 2.750 0.3314 0.09613 0.08070 -0.0665 0.9114 1.0012 3.000 0.2877 0.09582 0.08027 -0.0578 0.9647 1.0012 3.250 0.2516 0.09555 0.07989 -0.0491 0.9988 1.0012 3.500 0.2636 0.09870 0.08294 -0.0506 0.9988 1.0012 3.750 0.2756 0.10186 0.08599 -0.0521 0.9988 1.0012 4.000 0.2877 0.10504 0.08906 -0.0536 0.9988 1.0012 4.250 0.2998 0.10823 0.09216 -0.0550 0.9988 1.0012 4.500 0.3120 0.11144 0.09527 -0.0565 0.9988 1.0012 4.750 0.3241 0.11467 0.09840 -0.0580 0.9988 1.0012 5.000 0.3363 0.11791 0.10155 -0.0594 0.9988 1.0012