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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1817 0.03034 0.01618 -0.0312 1.0000 0.2243 -2.750 -0.1530 0.02811 0.01405 -0.0301 1.0000 0.2581 -2.500 -0.1281 0.02492 0.01202 -0.0285 1.0000 0.3612 -2.250 -0.0974 0.02154 0.00964 -0.0245 1.0000 1.0000 -2.000 -0.0717 0.02151 0.00879 -0.0233 1.0000 1.0000 -1.750 -0.0467 0.02151 0.00824 -0.0222 1.0000 1.0000 -1.500 -0.0221 0.02152 0.00779 -0.0213 1.0000 1.0000 -1.250 0.0023 0.02155 0.00743 -0.0204 1.0000 1.0000 -1.000 0.0266 0.02159 0.00716 -0.0195 1.0000 1.0000 -0.750 0.0508 0.02165 0.00695 -0.0186 1.0000 1.0000 -0.500 0.0750 0.02172 0.00689 -0.0177 1.0000 1.0000 -0.250 0.0993 0.02181 0.00696 -0.0169 1.0000 1.0000 0.000 0.1239 0.02191 0.00717 -0.0159 1.0000 1.0000 0.250 0.1493 0.02202 0.00758 -0.0149 1.0000 1.0000 0.500 0.1772 0.02208 0.00826 -0.0139 1.0000 1.0000 0.750 0.2655 0.02501 0.00958 -0.0205 0.4167 1.0000 1.000 0.2899 0.02680 0.01070 -0.0192 0.3783 1.0000 1.250 0.3186 0.02834 0.01192 -0.0186 0.3566 1.0000 1.500 0.3492 0.02989 0.01333 -0.0184 0.3399 1.0000 1.750 0.3800 0.03147 0.01487 -0.0184 0.3226 1.0000 2.000 0.4109 0.03316 0.01658 -0.0184 0.3077 1.0000 2.250 0.4442 0.03478 0.01858 -0.0190 0.3019 1.0000 2.500 0.4780 0.03664 0.02083 -0.0197 0.3013 1.0000 2.750 0.5113 0.03867 0.02338 -0.0207 0.3029 1.0000 3.000 0.5437 0.04096 0.02614 -0.0217 0.3063 1.0000 3.250 0.5753 0.04341 0.02911 -0.0230 0.3115 1.0000 3.500 0.6074 0.04621 0.03259 -0.0253 0.3209 1.0000 3.750 0.6389 0.04935 0.03631 -0.0281 0.3326 1.0000 4.000 0.6694 0.05293 0.04042 -0.0317 0.3482 1.0000 4.250 0.6984 0.05702 0.04505 -0.0369 0.3699 1.0000 4.500 0.7254 0.06153 0.04992 -0.0417 0.3931 1.0000 4.750 0.7450 0.06653 0.05535 -0.0504 0.4248 1.0000 5.000 0.7601 0.07195 0.06100 -0.0590 0.4609 1.0000