XFOIL Version 6.96 Calculated polar for: GOE 424 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.9788 0.07421 0.07086 -0.0889 1.0000 0.0134 -19.500 -0.9966 0.06830 0.06484 -0.0922 1.0000 0.0136 -19.250 -1.0100 0.06320 0.05964 -0.0950 1.0000 0.0137 -19.000 -1.0209 0.05859 0.05494 -0.0974 1.0000 0.0139 -18.750 -1.0283 0.05456 0.05082 -0.0995 1.0000 0.0141 -18.500 -1.0339 0.05090 0.04707 -0.1012 1.0000 0.0143 -18.250 -1.0375 0.04763 0.04373 -0.1025 1.0000 0.0145 -18.000 -1.0396 0.04468 0.04071 -0.1036 1.0000 0.0148 -17.750 -1.0412 0.04199 0.03795 -0.1043 1.0000 0.0150 -17.500 -1.0349 0.03942 0.03530 -0.1060 0.9967 0.0153 -17.250 -1.0249 0.03701 0.03281 -0.1080 0.9931 0.0156 -17.000 -1.0153 0.03484 0.03055 -0.1095 0.9880 0.0159 -16.750 -1.0035 0.03285 0.02848 -0.1109 0.9833 0.0161 -16.500 -0.9928 0.03101 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2.500 0.7441 0.00884 0.00301 -0.0785 0.4459 0.6527 2.750 0.7708 0.00884 0.00308 -0.0785 0.4428 0.6744 3.000 0.7967 0.00884 0.00315 -0.0783 0.4391 0.7004 3.250 0.8218 0.00888 0.00324 -0.0780 0.4351 0.7279 3.500 0.8483 0.00889 0.00332 -0.0780 0.4322 0.7505 3.750 0.8747 0.00890 0.00341 -0.0779 0.4294 0.7747 4.000 0.9003 0.00892 0.00351 -0.0776 0.4266 0.7999 4.250 0.9250 0.00895 0.00362 -0.0772 0.4238 0.8261 4.500 0.9488 0.00899 0.00374 -0.0766 0.4212 0.8526 4.750 0.9721 0.00907 0.00387 -0.0758 0.4185 0.8770 5.000 0.9971 0.00913 0.00399 -0.0755 0.4164 0.8962 5.250 1.0221 0.00920 0.00412 -0.0751 0.4139 0.9137 5.500 1.0468 0.00930 0.00426 -0.0747 0.4106 0.9314 5.750 1.0716 0.00943 0.00442 -0.0743 0.4067 0.9502 6.000 1.0989 0.00962 0.00460 -0.0746 0.4024 0.9683 6.250 1.1317 0.00976 0.00476 -0.0760 0.3981 0.9816 6.500 1.1644 0.00996 0.00494 -0.0776 0.3912 0.9914 6.750 1.1970 0.01019 0.00514 -0.0792 0.3836 0.9988 7.000 1.2169 0.01038 0.00531 -0.0781 0.3758 1.0000 7.250 1.2351 0.01060 0.00551 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12.250 1.5102 0.02526 0.01976 -0.0496 0.1985 1.0000 12.500 1.5184 0.02669 0.02122 -0.0485 0.1936 1.0000 12.750 1.5222 0.02852 0.02307 -0.0473 0.1883 1.0000 13.000 1.5277 0.03030 0.02487 -0.0464 0.1829 1.0000 13.250 1.5281 0.03257 0.02715 -0.0453 0.1756 1.0000 13.500 1.5277 0.03499 0.02960 -0.0442 0.1690 1.0000 13.750 1.5251 0.03769 0.03230 -0.0433 0.1612 1.0000 14.000 1.5216 0.04054 0.03517 -0.0424 0.1537 1.0000 14.250 1.5123 0.04397 0.03859 -0.0413 0.1448 1.0000 14.500 1.5025 0.04750 0.04212 -0.0404 0.1350 1.0000 14.750 1.4931 0.05103 0.04566 -0.0396 0.1266 1.0000 15.000 1.4830 0.05464 0.04927 -0.0388 0.1192 1.0000 15.250 1.4758 0.05803 0.05267 -0.0381 0.1126 1.0000 15.500 1.4702 0.06127 0.05592 -0.0376 0.1070 1.0000 15.750 1.4648 0.06456 0.05923 -0.0373 0.1017 1.0000 16.000 1.4589 0.06791 0.06259 -0.0369 0.0966 1.0000 16.250 1.4561 0.07097 0.06568 -0.0367 0.0910 1.0000 16.500 1.4494 0.07450 0.06921 -0.0365 0.0852 1.0000 16.750 1.4417 0.07816 0.07285 -0.0363 0.0756 1.0000 17.000 1.4142 0.08413 0.07870 -0.0360 0.0547 1.0000 17.250 1.3983 0.08893 0.08349 -0.0360 0.0452 1.0000