XFOIL Version 6.94 Calculated polar for: agora05 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0747 0.05813 0.03526 -0.0149 0.9992 1.0008 -2.750 -0.0568 0.05828 0.03475 -0.0149 0.9992 1.0008 -2.500 -0.0387 0.05854 0.03444 -0.0148 0.9992 1.0008 -2.250 -0.0203 0.05890 0.03429 -0.0146 0.9992 1.0008 -2.000 -0.0017 0.05934 0.03428 -0.0143 0.9992 1.0008 -1.750 0.0170 0.05984 0.03440 -0.0140 0.9992 1.0008 -1.500 0.0357 0.06042 0.03465 -0.0138 0.9992 1.0008 -1.250 0.0545 0.06106 0.03503 -0.0135 0.9992 1.0008 -1.000 0.0732 0.06177 0.03553 -0.0133 0.9992 1.0008 -0.750 0.0919 0.06257 0.03616 -0.0131 0.9992 1.0008 -0.500 0.1104 0.06346 0.03693 -0.0130 0.9992 1.0008 -0.250 0.1288 0.06445 0.03787 -0.0130 0.9992 1.0008 0.000 0.1467 0.06557 0.03898 -0.0131 0.9992 1.0008 0.250 0.1640 0.06684 0.04031 -0.0134 0.9992 1.0008 0.500 0.1802 0.06833 0.04189 -0.0139 0.9992 1.0008 0.750 0.1949 0.07010 0.04381 -0.0147 0.9992 1.0008 1.000 0.2071 0.07227 0.04615 -0.0159 0.9992 1.0008 1.250 0.2159 0.07498 0.04898 -0.0176 0.9992 1.0008 1.500 0.2217 0.07818 0.05222 -0.0197 0.9992 1.0008 1.750 0.2263 0.08160 0.05560 -0.0221 0.9992 1.0008 2.000 0.2317 0.08500 0.05891 -0.0244 0.9992 1.0008 2.250 0.2380 0.08832 0.06212 -0.0266 0.9992 1.0008 2.500 0.2452 0.09155 0.06524 -0.0286 0.9992 1.0008 2.750 0.2532 0.09473 0.06831 -0.0306 0.9992 1.0008 3.000 0.2620 0.09790 0.07136 -0.0326 0.9992 1.0008 3.250 0.2713 0.10102 0.07437 -0.0345 0.9992 1.0008 3.500 0.2811 0.10412 0.07735 -0.0363 0.9992 1.0008 3.750 0.2911 0.10720 0.08031 -0.0381 0.9992 1.0008 4.000 0.3014 0.11026 0.08326 -0.0399 0.9992 1.0008 4.250 0.3120 0.11331 0.08619 -0.0416 0.9992 1.0008 4.500 0.3229 0.11635 0.08912 -0.0433 0.9992 1.0008 4.750 0.3339 0.11939 0.09205 -0.0450 0.9992 1.0008 5.000 0.3451 0.12242 0.09497 -0.0467 0.9992 1.0008