XFOIL Version 6.94 Calculated polar for: nabad 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3684 0.05193 0.04026 -0.0029 0.9999 0.1699 -2.750 -0.3392 0.04858 0.03674 -0.0031 0.9999 0.1632 -2.500 -0.3058 0.04590 0.03329 -0.0035 0.9999 0.1550 -2.250 -0.2716 0.04319 0.03030 -0.0040 0.9999 0.1507 -2.000 -0.2347 0.04081 0.02750 -0.0046 0.9999 0.1471 -1.750 -0.1966 0.03871 0.02507 -0.0053 0.9999 0.1462 -1.500 -0.1588 0.03682 0.02305 -0.0059 0.9999 0.1484 -1.250 -0.1215 0.03519 0.02136 -0.0064 0.9999 0.1533 -1.000 -0.0862 0.03364 0.02007 -0.0067 0.9999 0.1620 -0.750 -0.0522 0.03225 0.01901 -0.0068 0.9999 0.1750 -0.500 -0.0195 0.03087 0.01816 -0.0067 0.9999 0.1963 -0.250 0.0142 0.02910 0.01747 -0.0071 0.9999 0.2494 0.000 0.1006 0.02654 0.01699 -0.0165 0.9999 1.0001 0.250 0.1573 0.02752 0.01686 -0.0188 0.6890 1.0001 0.500 0.1815 0.02841 0.01684 -0.0161 0.5978 1.0001 0.750 0.2088 0.02914 0.01697 -0.0150 0.5568 1.0001 1.000 0.2352 0.02988 0.01708 -0.0138 0.5389 1.0001 1.250 0.2635 0.03061 0.01736 -0.0132 0.5276 1.0001 1.500 0.2923 0.03135 0.01774 -0.0128 0.5202 1.0001 1.750 0.3220 0.03211 0.01820 -0.0128 0.5157 1.0001 2.000 0.3530 0.03286 0.01888 -0.0132 0.5128 1.0001 2.250 0.3837 0.03365 0.01970 -0.0139 0.5120 1.0001 2.500 0.4139 0.03446 0.02060 -0.0146 0.5104 1.0001 2.750 0.4350 0.03528 0.02106 -0.0126 0.4806 1.0001 3.000 0.4512 0.03630 0.02146 -0.0095 0.4341 1.0001 3.250 0.4713 0.03715 0.02194 -0.0076 0.4025 1.0001 3.500 0.4934 0.03769 0.02234 -0.0064 0.3779 1.0001 3.750 0.5152 0.03817 0.02266 -0.0051 0.3533 1.0001 4.000 0.5367 0.03859 0.02293 -0.0038 0.3274 1.0001 4.250 0.5597 0.03900 0.02331 -0.0028 0.3059 1.0001 4.500 0.5834 0.03958 0.02384 -0.0017 0.2883 1.0001 4.750 0.6072 0.04035 0.02458 -0.0005 0.2685 1.0001 5.000 0.6321 0.04144 0.02584 0.0004 0.2454 1.0001