XFOIL Version 6.94 Calculated polar for: GOE 123 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 0.4611 0.01130 0.00495 -0.0941 0.7802 0.0483 0.500 0.5166 0.01015 0.00361 -0.0935 0.7611 0.0504 1.000 0.5715 0.00972 0.00313 -0.0929 0.7348 0.0527 1.500 0.6263 0.00932 0.00264 -0.0922 0.7056 0.0551 2.000 0.6810 0.00909 0.00241 -0.0916 0.6694 0.0588 2.500 0.7350 0.00903 0.00227 -0.0908 0.6209 0.0705 3.000 0.7877 0.00923 0.00238 -0.0900 0.5602 0.1064 4.000 0.8871 0.00856 0.00305 -0.0877 0.4520 1.0000 4.500 0.9384 0.00921 0.00339 -0.0868 0.3962 1.0000 5.000 0.9888 0.00996 0.00383 -0.0859 0.3418 1.0000 5.500 1.0401 0.01060 0.00431 -0.0850 0.3085 1.0000 6.000 1.0911 0.01124 0.00482 -0.0842 0.2781 1.0000 6.500 1.1416 0.01189 0.00537 -0.0834 0.2483 1.0000 7.000 1.1911 0.01262 0.00599 -0.0824 0.2117 1.0000 7.500 1.2376 0.01365 0.00678 -0.0811 0.1625 1.0000 8.000 1.2759 0.01560 0.00817 -0.0788 0.0772 1.0000 8.500 1.3154 0.01732 0.00968 -0.0765 0.0425 1.0000 9.000 1.3563 0.01874 0.01112 -0.0743 0.0362 1.0000 9.500 1.3925 0.02042 0.01287 -0.0715 0.0330 1.0000 10.000 1.4255 0.02214 0.01473 -0.0684 0.0308 1.0000 10.500 1.4466 0.02435 0.01705 -0.0636 0.0289 1.0000 11.000 1.4651 0.02658 0.01944 -0.0588 0.0275 1.0000 11.500 1.4767 0.02955 0.02249 -0.0541 0.0264 1.0000 12.000 1.4884 0.03283 0.02596 -0.0501 0.0254 1.0000 12.500 1.5026 0.03605 0.02939 -0.0468 0.0241 1.0000 13.000 1.5134 0.03980 0.03324 -0.0438 0.0230 1.0000 13.500 1.5214 0.04412 0.03776 -0.0409 0.0220 1.0000 14.000 1.5271 0.04869 0.04263 -0.0389 0.0210 1.0000 14.500 1.5314 0.05358 0.04772 -0.0374 0.0202 1.0000 15.000 1.1559 0.05881 0.05404 -0.0131 0.0221 1.0000 15.500 1.1491 0.06384 0.05928 -0.0120 0.0219 1.0000 16.000 1.1322 0.07065 0.06639 -0.0138 0.0214 1.0000 17.500 1.4598 0.10418 0.09996 -0.0498 0.0155 1.0000 18.500 1.0335 0.11098 0.10792 -0.0344 0.0178 1.0000