XFOIL Version 6.94 Calculated polar for: GOA 1b 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 -0.1723 0.02512 0.01324 -0.0308 1.0000 0.0962 -2.750 -0.1448 0.02352 0.01152 -0.0297 1.0000 0.1020 -2.500 -0.1181 0.02225 0.01013 -0.0285 1.0000 0.1121 -2.250 -0.0918 0.02094 0.00883 -0.0275 1.0000 0.1300 -1.500 -0.0183 0.01612 0.00652 -0.0225 1.0000 1.0000 -1.250 0.0066 0.01618 0.00631 -0.0215 1.0000 1.0000 -1.000 0.0314 0.01624 0.00614 -0.0205 1.0000 1.0000 -0.750 0.0562 0.01631 0.00602 -0.0196 1.0000 1.0000 -0.500 0.0812 0.01638 0.00608 -0.0186 1.0000 1.0000 -0.250 0.1071 0.01646 0.00635 -0.0175 1.0000 1.0000 0.000 0.1836 0.01961 0.00672 -0.0236 0.2712 1.0000 0.250 0.2075 0.02073 0.00732 -0.0226 0.2430 1.0000 0.500 0.2329 0.02166 0.00796 -0.0218 0.2215 1.0000 0.750 0.2595 0.02271 0.00875 -0.0211 0.2104 1.0000 1.000 0.2888 0.02366 0.00968 -0.0205 0.2053 1.0000 1.250 0.3197 0.02471 0.01075 -0.0203 0.2019 1.0000 1.500 0.3517 0.02587 0.01202 -0.0202 0.2002 1.0000 1.750 0.3832 0.02716 0.01348 -0.0201 0.2004 1.0000 2.000 0.4141 0.02862 0.01517 -0.0200 0.2022 1.0000 2.250 0.4441 0.03030 0.01711 -0.0199 0.2052 1.0000 2.500 0.4734 0.03229 0.01929 -0.0199 0.2088 1.0000 2.750 0.5053 0.03404 0.02169 -0.0201 0.2172 1.0000 3.000 0.5344 0.03660 0.02452 -0.0203 0.2254 1.0000 3.250 0.5662 0.03890 0.02750 -0.0212 0.2363 1.0000 3.500 0.5938 0.04167 0.03064 -0.0217 0.2397 1.0000 3.750 0.6245 0.04498 0.03446 -0.0232 0.2576 1.0000 4.000 0.6751 0.04981 0.04058 -0.0323 0.3456 1.0000 4.250 0.6588 0.06308 0.05499 -0.0876 0.6879 1.0000 4.500 0.5677 0.06574 0.05730 -0.0868 0.8350 1.0000 5.000 0.4382 0.06308 0.05422 -0.0625 1.0000 1.0000