XFOIL Version 6.94 Calculated polar for: GOA 1 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 -0.1832 0.03217 0.01720 -0.0319 1.0000 0.1661 -2.750 -0.1545 0.03025 0.01513 -0.0309 1.0000 0.1809 -2.500 -0.1276 0.02829 0.01332 -0.0298 1.0000 0.2074 -2.250 -0.1017 0.02612 0.01166 -0.0287 1.0000 0.2607 -2.000 -0.0766 0.02151 0.00957 -0.0248 1.0000 1.0000 -1.750 -0.0499 0.02150 0.00856 -0.0233 1.0000 1.0000 -1.500 -0.0245 0.02151 0.00804 -0.0221 1.0000 1.0000 -1.250 0.0004 0.02155 0.00772 -0.0210 1.0000 1.0000 -1.000 0.0251 0.02160 0.00750 -0.0200 1.0000 1.0000 -0.750 0.0496 0.02166 0.00732 -0.0190 1.0000 1.0000 -0.500 0.0742 0.02174 0.00724 -0.0180 1.0000 1.0000 -0.250 0.0988 0.02183 0.00724 -0.0170 1.0000 1.0000 0.000 0.1240 0.02192 0.00720 -0.0160 1.0000 1.0000 0.250 0.1495 0.02202 0.00759 -0.0149 1.0000 1.0000 0.500 0.1775 0.02208 0.00826 -0.0139 1.0000 1.0000 0.750 0.2657 0.02502 0.00959 -0.0204 0.4163 1.0000 1.000 0.2901 0.02681 0.01071 -0.0192 0.3782 1.0000 1.250 0.3188 0.02835 0.01195 -0.0186 0.3566 1.0000 1.500 0.3494 0.02989 0.01336 -0.0184 0.3400 1.0000 1.750 0.3801 0.03148 0.01490 -0.0183 0.3226 1.0000 2.000 0.4110 0.03317 0.01661 -0.0185 0.3077 1.0000 2.250 0.4443 0.03480 0.01868 -0.0190 0.3020 1.0000 2.500 0.4780 0.03665 0.02091 -0.0198 0.3013 1.0000 2.750 0.5113 0.03868 0.02340 -0.0208 0.3029 1.0000 3.000 0.5436 0.04097 0.02613 -0.0218 0.3063 1.0000 3.250 0.5751 0.04342 0.02908 -0.0231 0.3116 1.0000 3.500 0.6071 0.04623 0.03248 -0.0254 0.3210 1.0000 3.750 0.6386 0.04936 0.03618 -0.0282 0.3326 1.0000 4.000 0.6690 0.05293 0.04026 -0.0318 0.3482 1.0000 4.250 0.6981 0.05703 0.04488 -0.0370 0.3699 1.0000 4.500 0.7251 0.06154 0.04973 -0.0419 0.3931 1.0000 4.750 0.7447 0.06654 0.05516 -0.0505 0.4249 1.0000 5.000 0.7600 0.07196 0.06080 -0.0591 0.4611 1.0000