XFOIL Version 6.94 Calculated polar for: GOE 503 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6118 0.01140 0.00367 -0.1091 0.5669 0.1856 1.000 0.7022 0.01155 0.00400 -0.1051 0.5468 0.2803 1.500 0.7449 0.01159 0.00409 -0.1026 0.5373 0.3052 2.000 0.7856 0.01161 0.00420 -0.0997 0.5281 0.3451 2.500 0.8207 0.01138 0.00434 -0.0958 0.5194 0.4780 3.000 1.0212 0.01157 0.00556 -0.1275 0.5065 0.9936 3.500 1.0983 0.01191 0.00583 -0.1327 0.4968 1.0000 4.000 1.1272 0.01216 0.00603 -0.1277 0.4884 1.0000 4.500 1.1426 0.01233 0.00618 -0.1197 0.4778 1.0000 5.000 1.1459 0.01252 0.00629 -0.1093 0.4603 1.0000 5.500 1.1609 0.01284 0.00657 -0.1015 0.4410 1.0000 6.000 1.1853 0.01330 0.00701 -0.0960 0.4256 1.0000 6.500 1.1919 0.01427 0.00775 -0.0875 0.3825 1.0000 7.000 1.2015 0.01565 0.00888 -0.0803 0.3396 1.0000 7.500 1.2029 0.01774 0.01064 -0.0727 0.2861 1.0000 8.000 1.2012 0.02041 0.01299 -0.0655 0.2308 1.0000 8.500 1.1704 0.02533 0.01741 -0.0563 0.1479 1.0000 9.000 1.1075 0.03381 0.02533 -0.0467 0.0050 1.0000 9.500 1.1293 0.03639 0.02798 -0.0448 0.0030 1.0000 10.000 1.1486 0.03928 0.03095 -0.0431 0.0030 1.0000 10.500 1.1653 0.04251 0.03435 -0.0414 0.0030 1.0000 11.000 1.1790 0.04618 0.03818 -0.0397 0.0029 1.0000 11.500 1.1911 0.05011 0.04228 -0.0382 0.0030 1.0000 12.000 1.2005 0.05446 0.04680 -0.0369 0.0031 1.0000 12.500 1.2049 0.05950 0.05205 -0.0358 0.0032 1.0000 13.000 1.2052 0.06519 0.05796 -0.0348 0.0035 1.0000 13.500 1.2061 0.07093 0.06388 -0.0342 0.0036 1.0000 14.000 1.2034 0.07729 0.07045 -0.0338 0.0038 1.0000 14.500 1.2012 0.08377 0.07711 -0.0336 0.0045 1.0000 15.000 1.1972 0.09062 0.08414 -0.0338 0.0048 1.0000 15.500 1.1903 0.09805 0.09179 -0.0342 0.0057 1.0000 16.000 1.1807 0.10595 0.09989 -0.0351 0.0058 1.0000 16.500 1.1692 0.11429 0.10842 -0.0364 0.0062 1.0000 17.000 1.1551 0.12315 0.11747 -0.0380 0.0067 1.0000