XFOIL Version 6.94 Calculated polar for: AG45c -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 0.0754 0.00610 0.00177 -0.0273 0.6935 0.8693 0.500 0.1406 0.00595 0.00157 -0.0285 0.6675 1.0000 1.000 0.1968 0.00602 0.00142 -0.0282 0.6396 1.0000 1.500 0.2528 0.00614 0.00132 -0.0279 0.6078 1.0000 2.000 0.3088 0.00631 0.00130 -0.0277 0.5728 1.0000 2.500 0.3646 0.00654 0.00133 -0.0274 0.5374 1.0000 3.000 0.4204 0.00679 0.00141 -0.0272 0.5003 1.0000 3.500 0.4762 0.00707 0.00156 -0.0271 0.4634 1.0000 4.000 0.5319 0.00740 0.00174 -0.0269 0.4253 1.0000 4.500 0.5873 0.00776 0.00199 -0.0268 0.3839 1.0000 5.000 0.6426 0.00819 0.00229 -0.0266 0.3397 1.0000 5.500 0.6974 0.00871 0.00269 -0.0265 0.2927 1.0000 6.000 0.7518 0.00932 0.00316 -0.0264 0.2437 1.0000 6.500 0.8057 0.01006 0.00377 -0.0263 0.1952 1.0000 7.000 0.8591 0.01087 0.00446 -0.0261 0.1516 1.0000 7.500 0.9119 0.01178 0.00529 -0.0259 0.1139 1.0000 8.000 0.9642 0.01274 0.00621 -0.0256 0.0811 1.0000 8.500 1.0151 0.01398 0.00738 -0.0253 0.0523 1.0000 9.000 1.0647 0.01550 0.00899 -0.0247 0.0326 1.0000 10.500 1.1971 0.02244 0.01663 -0.0215 0.0135 1.0000 11.000 1.2214 0.02783 0.02247 -0.0193 0.0108 1.0000 11.500 1.2538 0.03110 0.02618 -0.0177 0.0100 1.0000 12.000 1.2715 0.03594 0.03151 -0.0158 0.0092 1.0000 12.500 1.2668 0.04190 0.03796 -0.0139 0.0088 1.0000 13.000 1.2473 0.05116 0.04767 -0.0185 0.0086 1.0000 13.500 1.2196 0.06353 0.06044 -0.0265 0.0086 1.0000 14.000 1.1828 0.07875 0.07601 -0.0364 0.0087 1.0000 14.500 1.1384 0.09661 0.09416 -0.0472 0.0089 1.0000 15.000 1.0875 0.11760 0.11537 -0.0591 0.0094 1.0000 15.500 1.0135 0.14775 0.14563 -0.0747 0.0104 1.0000