XFOIL Version 6.94 Calculated polar for: GOE 363 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.6713 0.01002 0.00265 -0.1126 0.5258 0.1708 1.000 0.7236 0.01016 0.00281 -0.1119 0.4905 0.2255 1.500 0.7736 0.00885 0.00314 -0.1111 0.4466 1.0000 2.000 0.8211 0.00960 0.00339 -0.1096 0.3714 1.0000 2.500 0.8687 0.01040 0.00379 -0.1082 0.3202 1.0000 3.000 0.9187 0.01101 0.00416 -0.1073 0.2949 1.0000 3.500 0.9687 0.01159 0.00457 -0.1064 0.2788 1.0000 4.000 1.0183 0.01218 0.00497 -0.1055 0.2594 1.0000 4.500 1.0656 0.01291 0.00538 -0.1044 0.2128 1.0000 5.000 1.1111 0.01379 0.00590 -0.1030 0.1661 1.0000 5.500 1.1462 0.01553 0.00710 -0.1001 0.0980 1.0000 6.000 1.1778 0.01749 0.00867 -0.0965 0.0052 1.0000 6.500 1.2220 0.01825 0.00947 -0.0948 0.0051 1.0000 7.000 1.2629 0.01906 0.01036 -0.0925 0.0057 1.0000 7.500 1.3012 0.02000 0.01141 -0.0898 0.0065 1.0000 8.000 1.3377 0.02107 0.01261 -0.0869 0.0076 1.0000 8.500 1.3727 0.02224 0.01392 -0.0840 0.0091 1.0000 9.000 1.4037 0.02374 0.01560 -0.0808 0.0101 1.0000 9.500 1.4307 0.02560 0.01769 -0.0773 0.0114 1.0000 10.500 1.4639 0.03127 0.02382 -0.0696 0.0139 1.0000 11.000 1.4809 0.03434 0.02712 -0.0665 0.0160 1.0000 11.500 1.4700 0.04020 0.03325 -0.0629 0.0165 1.0000 12.000 1.4659 0.04595 0.03923 -0.0607 0.0168 1.0000 12.500 1.4539 0.05318 0.04672 -0.0596 0.0170 1.0000 13.000 1.4353 0.06211 0.05590 -0.0600 0.0171 1.0000 13.500 1.4159 0.07181 0.06584 -0.0613 0.0172 1.0000 14.000 1.3965 0.08150 0.07570 -0.0627 0.0171 1.0000 14.500 1.3865 0.08895 0.08317 -0.0624 0.0167 1.0000 15.000 1.3862 0.09533 0.08968 -0.0622 0.0160 1.0000 15.500 1.4027 0.09772 0.09198 -0.0587 0.0136 1.0000