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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1733 0.02642 0.01394 -0.0310 1.0000 0.1155 -2.750 -0.1458 0.02470 0.01217 -0.0299 1.0000 0.1227 -2.500 -0.1191 0.02323 0.01061 -0.0285 1.0000 0.1341 -2.250 -0.0928 0.02168 0.00917 -0.0275 1.0000 0.1645 -2.000 -0.0700 0.01714 0.00777 -0.0247 1.0000 1.0000 -1.750 -0.0442 0.01715 0.00714 -0.0234 1.0000 1.0000 -1.500 -0.0191 0.01718 0.00680 -0.0222 1.0000 1.0000 -1.250 0.0057 0.01724 0.00654 -0.0212 1.0000 1.0000 -1.000 0.0303 0.01729 0.00633 -0.0203 1.0000 1.0000 -0.750 0.0549 0.01736 0.00619 -0.0193 1.0000 1.0000 -0.500 0.0797 0.01744 0.00621 -0.0184 1.0000 1.0000 -0.250 0.1049 0.01752 0.00641 -0.0174 1.0000 1.0000 0.000 0.1321 0.01758 0.00687 -0.0163 1.0000 1.0000 0.250 0.2085 0.02117 0.00750 -0.0224 0.2869 1.0000 0.500 0.2329 0.02232 0.00821 -0.0213 0.2619 1.0000 0.750 0.2587 0.02345 0.00900 -0.0205 0.2408 1.0000 1.000 0.2864 0.02456 0.00992 -0.0199 0.2278 1.0000 1.250 0.3170 0.02567 0.01102 -0.0196 0.2229 1.0000 1.500 0.3489 0.02684 0.01232 -0.0195 0.2199 1.0000 1.750 0.3817 0.02813 0.01380 -0.0196 0.2187 1.0000 2.000 0.4135 0.02955 0.01549 -0.0195 0.2194 1.0000 2.250 0.4449 0.03120 0.01743 -0.0196 0.2216 1.0000 2.500 0.4750 0.03311 0.01959 -0.0197 0.2247 1.0000 2.750 0.5062 0.03496 0.02198 -0.0200 0.2307 1.0000 3.000 0.5371 0.03733 0.02480 -0.0205 0.2395 1.0000 3.250 0.5692 0.03984 0.02792 -0.0216 0.2525 1.0000 3.500 0.5968 0.04294 0.03124 -0.0220 0.2593 1.0000 3.750 0.6279 0.04575 0.03479 -0.0242 0.2713 1.0000 4.000 0.6581 0.04943 0.03893 -0.0264 0.2907 1.0000 4.250 0.7054 0.05486 0.04545 -0.0378 0.3755 1.0000 4.750 0.6324 0.07036 0.06147 -0.0884 0.7371 1.0000 5.000 0.5115 0.06887 0.05956 -0.0779 0.9333 1.0000