XFOIL Version 6.94 Calculated polar for: GOA 1 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 -0.1780 0.02664 0.01483 -0.0321 1.0000 0.0909 -2.750 -0.1500 0.02499 0.01289 -0.0308 1.0000 0.0924 -2.500 -0.1220 0.02353 0.01115 -0.0295 1.0000 0.0948 -2.250 -0.0952 0.02228 0.00975 -0.0281 1.0000 0.0982 -2.000 -0.0690 0.02105 0.00856 -0.0271 1.0000 0.1063 -1.750 -0.0420 0.01991 0.00755 -0.0263 1.0000 0.1299 -1.500 -0.0215 0.01611 0.00681 -0.0236 1.0000 1.0000 -1.250 0.0042 0.01615 0.00644 -0.0224 1.0000 1.0000 -1.000 0.0295 0.01622 0.00631 -0.0212 1.0000 1.0000 -0.750 0.0548 0.01630 0.00628 -0.0201 1.0000 1.0000 -0.500 0.0803 0.01639 0.00637 -0.0189 1.0000 1.0000 -0.250 0.1067 0.01646 0.00660 -0.0177 1.0000 1.0000 0.000 0.1837 0.01962 0.00670 -0.0236 0.2706 1.0000 0.250 0.2077 0.02073 0.00730 -0.0226 0.2427 1.0000 0.500 0.2331 0.02166 0.00795 -0.0218 0.2213 1.0000 0.750 0.2597 0.02272 0.00875 -0.0210 0.2104 1.0000 1.000 0.2890 0.02367 0.00967 -0.0205 0.2053 1.0000 1.250 0.3199 0.02472 0.01077 -0.0203 0.2019 1.0000 1.500 0.3519 0.02588 0.01204 -0.0202 0.2003 1.0000 1.750 0.3834 0.02717 0.01350 -0.0201 0.2005 1.0000 2.000 0.4142 0.02863 0.01520 -0.0200 0.2023 1.0000 2.250 0.4441 0.03031 0.01718 -0.0199 0.2053 1.0000 2.500 0.4733 0.03230 0.01934 -0.0199 0.2089 1.0000 2.750 0.5052 0.03406 0.02170 -0.0202 0.2174 1.0000 3.000 0.5343 0.03662 0.02450 -0.0204 0.2255 1.0000 3.250 0.5660 0.03891 0.02746 -0.0213 0.2364 1.0000 3.500 0.5936 0.04167 0.03053 -0.0218 0.2397 1.0000 3.750 0.6242 0.04498 0.03433 -0.0234 0.2575 1.0000 4.000 0.6749 0.04981 0.04044 -0.0326 0.3459 1.0000 4.250 0.6581 0.06309 0.05497 -0.0877 0.6883 1.0000 4.500 0.5668 0.06571 0.05730 -0.0868 0.8354 1.0000 5.000 0.4381 0.06311 0.05431 -0.0625 1.0000 1.0000