XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1597 0.02469 0.01297 -0.0318 1.0000 0.1003 -2.750 -0.1331 0.02336 0.01157 -0.0306 1.0000 0.1094 -2.500 -0.1068 0.02206 0.01028 -0.0295 1.0000 0.1233 -2.250 -0.0800 0.02051 0.00903 -0.0286 1.0000 0.1662 -2.000 -0.0594 0.01662 0.00808 -0.0256 1.0000 1.0000 -1.750 -0.0341 0.01667 0.00769 -0.0243 1.0000 1.0000 -1.500 -0.0092 0.01674 0.00752 -0.0232 1.0000 1.0000 -1.250 0.0157 0.01683 0.00743 -0.0221 1.0000 1.0000 -1.000 0.0407 0.01691 0.00740 -0.0210 1.0000 1.0000 -0.750 0.0664 0.01699 0.00752 -0.0199 1.0000 1.0000 -0.500 0.1541 0.01971 0.00740 -0.0276 0.2695 1.0000 -0.250 0.1777 0.02075 0.00791 -0.0265 0.2439 1.0000 0.000 0.2024 0.02167 0.00845 -0.0256 0.2227 1.0000 0.250 0.2282 0.02263 0.00908 -0.0248 0.2079 1.0000 0.500 0.2567 0.02351 0.00984 -0.0241 0.2020 1.0000 0.750 0.2869 0.02452 0.01074 -0.0237 0.1981 1.0000 1.000 0.3188 0.02564 0.01182 -0.0237 0.1960 1.0000 1.250 0.3511 0.02683 0.01307 -0.0237 0.1956 1.0000 1.500 0.3829 0.02808 0.01450 -0.0236 0.1968 1.0000 1.750 0.4138 0.02948 0.01616 -0.0234 0.1997 1.0000 2.000 0.4439 0.03114 0.01806 -0.0234 0.2038 1.0000 2.250 0.4732 0.03309 0.02019 -0.0234 0.2081 1.0000 2.500 0.5049 0.03481 0.02246 -0.0236 0.2170 1.0000 2.750 0.5338 0.03731 0.02516 -0.0238 0.2254 1.0000 3.000 0.5655 0.03975 0.02819 -0.0245 0.2412 1.0000 3.250 0.6005 0.04246 0.03172 -0.0265 0.2686 1.0000 3.500 0.6316 0.04578 0.03551 -0.0284 0.2899 1.0000 3.750 0.6709 0.04977 0.04029 -0.0341 0.3409 1.0000 4.000 0.6419 0.06354 0.05537 -0.0887 0.7012 1.0000 4.250 0.5525 0.06650 0.05807 -0.0868 0.8343 1.0000 4.750 0.4197 0.06395 0.05513 -0.0621 1.0000 1.0000 5.000 0.4313 0.06693 0.05807 -0.0634 1.0000 1.0000