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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1678 0.04647 0.01933 -0.0224 1.0000 1.0000 -2.750 -0.1435 0.04626 0.01802 -0.0220 1.0000 1.0000 -2.500 -0.1194 0.04609 0.01689 -0.0214 1.0000 1.0000 -2.250 -0.0954 0.04595 0.01592 -0.0208 1.0000 1.0000 -2.000 -0.0716 0.04585 0.01497 -0.0202 1.0000 1.0000 -1.750 -0.0478 0.04578 0.01427 -0.0195 1.0000 1.0000 -1.500 -0.0242 0.04575 0.01370 -0.0188 1.0000 1.0000 -1.250 -0.0006 0.04574 0.01325 -0.0181 1.0000 1.0000 -1.000 0.0228 0.04576 0.01292 -0.0173 1.0000 1.0000 -0.750 0.0463 0.04581 0.01270 -0.0166 1.0000 1.0000 -0.500 0.0697 0.04589 0.01252 -0.0158 1.0000 1.0000 -0.250 0.0931 0.04600 0.01253 -0.0151 1.0000 1.0000 0.000 0.1164 0.04613 0.01267 -0.0143 1.0000 1.0000 0.250 0.1398 0.04628 0.01293 -0.0135 1.0000 1.0000 0.500 0.1632 0.04647 0.01332 -0.0128 1.0000 1.0000 0.750 0.1868 0.04668 0.01384 -0.0120 1.0000 1.0000 1.000 0.2105 0.04691 0.01446 -0.0113 1.0000 1.0000 1.250 0.2346 0.04717 0.01528 -0.0106 1.0000 1.0000 1.500 0.2594 0.04745 0.01630 -0.0100 1.0000 1.0000 1.750 0.2851 0.04776 0.01756 -0.0096 1.0000 1.0000 2.000 0.3124 0.04813 0.01914 -0.0097 1.0000 1.0000 2.250 0.3403 0.04870 0.02120 -0.0109 1.0000 1.0000 2.500 0.3620 0.04993 0.02383 -0.0137 1.0000 1.0000 2.750 0.3680 0.05235 0.02681 -0.0168 1.0000 1.0000 3.000 0.3664 0.05548 0.02977 -0.0196 1.0000 1.0000 3.250 0.3666 0.05861 0.03263 -0.0223 1.0000 1.0000 3.500 0.3696 0.06163 0.03543 -0.0248 1.0000 1.0000 3.750 0.3744 0.06458 0.03819 -0.0271 1.0000 1.0000 4.000 0.3803 0.06748 0.04094 -0.0293 1.0000 1.0000 4.250 0.3870 0.07036 0.04370 -0.0314 1.0000 1.0000 4.500 0.3943 0.07324 0.04649 -0.0334 1.0000 1.0000 4.750 0.4022 0.07613 0.04929 -0.0353 1.0000 1.0000 5.000 0.4104 0.07904 0.05212 -0.0370 1.0000 1.0000