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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1818 0.03053 0.01654 -0.0319 1.0000 0.1374 -2.750 -0.1525 0.02872 0.01445 -0.0308 1.0000 0.1419 -2.500 -0.1247 0.02706 0.01259 -0.0294 1.0000 0.1508 -2.250 -0.0984 0.02540 0.01107 -0.0283 1.0000 0.1714 -2.000 -0.0725 0.02312 0.00963 -0.0275 1.0000 0.2468 -1.750 -0.0492 0.01981 0.00822 -0.0238 1.0000 1.0000 -1.500 -0.0236 0.01982 0.00764 -0.0225 1.0000 1.0000 -1.250 0.0015 0.01986 0.00733 -0.0214 1.0000 1.0000 -1.000 0.0263 0.01992 0.00714 -0.0203 1.0000 1.0000 -0.750 0.0510 0.01999 0.00699 -0.0193 1.0000 1.0000 -0.500 0.0757 0.02007 0.00694 -0.0183 1.0000 1.0000 -0.250 0.1006 0.02016 0.00698 -0.0173 1.0000 1.0000 0.000 0.1263 0.02025 0.00701 -0.0161 1.0000 1.0000 0.250 0.1531 0.02032 0.00751 -0.0150 1.0000 1.0000 0.500 0.2368 0.02321 0.00853 -0.0212 0.3749 1.0000 0.750 0.2607 0.02476 0.00947 -0.0199 0.3391 1.0000 1.000 0.2880 0.02611 0.01052 -0.0191 0.3183 1.0000 1.250 0.3164 0.02748 0.01167 -0.0186 0.2989 1.0000 1.500 0.3459 0.02889 0.01295 -0.0183 0.2810 1.0000 1.750 0.3775 0.03035 0.01446 -0.0183 0.2714 1.0000 2.000 0.4109 0.03185 0.01622 -0.0186 0.2686 1.0000 2.250 0.4443 0.03348 0.01827 -0.0191 0.2679 1.0000 2.500 0.4774 0.03528 0.02047 -0.0197 0.2693 1.0000 2.750 0.5092 0.03731 0.02289 -0.0202 0.2724 1.0000 3.000 0.5402 0.03961 0.02555 -0.0209 0.2766 1.0000 3.250 0.5731 0.04201 0.02862 -0.0225 0.2861 1.0000 3.500 0.6043 0.04481 0.03191 -0.0241 0.2963 1.0000 3.750 0.6345 0.04820 0.03564 -0.0259 0.3092 1.0000 4.000 0.6677 0.05163 0.03987 -0.0306 0.3325 1.0000 4.250 0.6962 0.05565 0.04438 -0.0352 0.3526 1.0000 4.500 0.7223 0.06007 0.04917 -0.0401 0.3740 1.0000 5.000 0.7487 0.07249 0.06237 -0.0690 0.5057 1.0000