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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1661 0.03670 0.01699 -0.0243 1.0000 1.0000 -2.750 -0.1409 0.03654 0.01571 -0.0237 1.0000 1.0000 -2.500 -0.1163 0.03641 0.01466 -0.0230 1.0000 1.0000 -2.250 -0.0920 0.03631 0.01376 -0.0223 1.0000 1.0000 -2.000 -0.0679 0.03624 0.01293 -0.0215 1.0000 1.0000 -1.750 -0.0441 0.03620 0.01230 -0.0207 1.0000 1.0000 -1.500 -0.0203 0.03619 0.01179 -0.0199 1.0000 1.0000 -1.250 0.0034 0.03620 0.01140 -0.0191 1.0000 1.0000 -1.000 0.0270 0.03624 0.01112 -0.0183 1.0000 1.0000 -0.750 0.0506 0.03630 0.01095 -0.0175 1.0000 1.0000 -0.500 0.0742 0.03639 0.01082 -0.0167 1.0000 1.0000 -0.250 0.0977 0.03649 0.01087 -0.0158 1.0000 1.0000 0.000 0.1214 0.03663 0.01104 -0.0150 1.0000 1.0000 0.250 0.1451 0.03678 0.01134 -0.0141 1.0000 1.0000 0.500 0.1691 0.03695 0.01177 -0.0132 1.0000 1.0000 0.750 0.1936 0.03713 0.01236 -0.0124 1.0000 1.0000 1.000 0.2191 0.03732 0.01310 -0.0116 1.0000 1.0000 1.250 0.2468 0.03749 0.01413 -0.0110 1.0000 1.0000 1.500 0.2786 0.03765 0.01555 -0.0112 1.0000 1.0000 1.750 0.3127 0.03808 0.01755 -0.0136 1.0000 1.0000 2.000 0.3294 0.03978 0.02030 -0.0175 1.0000 1.0000 2.250 0.3264 0.04288 0.02335 -0.0207 1.0000 1.0000 2.750 0.5172 0.05422 0.03499 -0.0590 0.7599 1.0000 3.000 0.5530 0.05730 0.03806 -0.0622 0.7232 1.0000 3.250 0.5762 0.06061 0.04134 -0.0647 0.7066 1.0000 3.500 0.5972 0.06414 0.04482 -0.0672 0.6974 1.0000 3.750 0.6047 0.06765 0.04828 -0.0686 0.6946 1.0000 4.000 0.6115 0.07120 0.05175 -0.0701 0.6940 1.0000 4.250 0.6168 0.07474 0.05520 -0.0713 0.6955 1.0000 4.500 0.6240 0.07840 0.05878 -0.0728 0.6984 1.0000 4.750 0.6190 0.08166 0.06192 -0.0730 0.7046 1.0000 5.000 0.6222 0.08525 0.06541 -0.0743 0.7113 1.0000