XFOIL Version 6.94 Calculated polar for: AG46c -03f 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0490 0.00532 0.00148 -0.0232 0.7648 0.9124 0.500 0.1156 0.00549 0.00132 -0.0250 0.7312 1.0000 1.000 0.1706 0.00575 0.00131 -0.0245 0.7049 1.0000 1.500 0.2264 0.00589 0.00129 -0.0242 0.6769 1.0000 2.000 0.2822 0.00597 0.00124 -0.0239 0.6427 1.0000 2.500 0.3377 0.00612 0.00123 -0.0235 0.5985 1.0000 3.000 0.3928 0.00639 0.00128 -0.0231 0.5440 1.0000 3.500 0.4477 0.00675 0.00141 -0.0228 0.4826 1.0000 4.000 0.5025 0.00722 0.00164 -0.0225 0.4141 1.0000 4.500 0.5571 0.00777 0.00193 -0.0224 0.3421 1.0000 5.000 0.6114 0.00843 0.00232 -0.0222 0.2689 1.0000 5.500 0.6655 0.00917 0.00284 -0.0221 0.2023 1.0000 6.000 0.7192 0.00996 0.00344 -0.0220 0.1506 1.0000 6.500 0.7728 0.01081 0.00418 -0.0218 0.1122 1.0000 7.000 0.8261 0.01163 0.00498 -0.0216 0.0828 1.0000 7.500 0.8788 0.01262 0.00598 -0.0212 0.0582 1.0000 8.000 0.9305 0.01383 0.00720 -0.0208 0.0391 1.0000 8.500 0.9793 0.01580 0.00932 -0.0201 0.0260 1.0000 9.000 1.0298 0.01710 0.01084 -0.0195 0.0198 1.0000 10.000 1.1170 0.02293 0.01739 -0.0170 0.0133 1.0000 10.500 1.1483 0.02807 0.02302 -0.0154 0.0107 1.0000 11.000 1.1745 0.03359 0.02924 -0.0138 0.0100 1.0000 11.500 1.1868 0.04008 0.03639 -0.0125 0.0095 1.0000 12.000 1.1693 0.04824 0.04518 -0.0124 0.0092 1.0000 12.500 1.1366 0.06125 0.05860 -0.0229 0.0093 1.0000 13.000 1.0950 0.08022 0.07789 -0.0381 0.0096 1.0000 13.500 1.0264 0.10655 0.10440 -0.0542 0.0103 1.0000