XFOIL Version 6.94 Calculated polar for: GOE 401 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 0.5007 0.01130 0.00382 -0.0796 0.5342 0.0433 1.000 0.5495 0.01087 0.00314 -0.0778 0.4929 0.0419 1.500 0.5987 0.01058 0.00268 -0.0761 0.4671 0.0411 2.000 0.6492 0.01050 0.00249 -0.0747 0.4481 0.0413 2.500 0.7002 0.01059 0.00245 -0.0734 0.4326 0.0440 3.000 0.7506 0.01044 0.00272 -0.0722 0.4186 0.2105 4.000 0.8900 0.00954 0.00317 -0.0790 0.3618 1.0000 4.500 0.9384 0.00991 0.00333 -0.0774 0.3215 1.0000 5.000 0.9844 0.01052 0.00363 -0.0755 0.2668 1.0000 5.500 1.0296 0.01124 0.00414 -0.0736 0.2212 1.0000 6.000 1.0543 0.01395 0.00576 -0.0688 0.0415 1.0000 6.500 1.0946 0.01518 0.00671 -0.0661 0.0062 1.0000 7.000 1.1405 0.01580 0.00742 -0.0643 0.0049 1.0000 7.500 1.1841 0.01660 0.00839 -0.0621 0.0047 1.0000 8.000 1.2262 0.01748 0.00943 -0.0598 0.0047 1.0000 8.500 1.2661 0.01847 0.01060 -0.0571 0.0048 1.0000 9.000 1.3026 0.01965 0.01197 -0.0539 0.0049 1.0000 9.500 1.3343 0.02103 0.01356 -0.0501 0.0051 1.0000 10.000 1.3574 0.02264 0.01538 -0.0450 0.0053 1.0000 10.500 1.3717 0.02458 0.01760 -0.0390 0.0054 1.0000 11.000 1.3756 0.02744 0.02074 -0.0330 0.0057 1.0000 11.500 1.3752 0.03118 0.02476 -0.0285 0.0059 1.0000 12.000 1.3690 0.03629 0.03016 -0.0257 0.0061 1.0000 12.500 1.3585 0.04272 0.03687 -0.0244 0.0064 1.0000 13.000 1.3473 0.04981 0.04424 -0.0242 0.0065 1.0000 13.500 1.3391 0.05707 0.05175 -0.0252 0.0058 1.0000 14.000 1.3256 0.06511 0.06008 -0.0257 0.0064 1.0000 14.500 1.3143 0.07349 0.06874 -0.0275 0.0061 1.0000 15.000 1.2969 0.08299 0.07856 -0.0292 0.0066 1.0000 15.500 1.2797 0.09332 0.08920 -0.0327 0.0066 1.0000 16.000 1.2575 0.10501 0.10121 -0.0373 0.0067 1.0000 16.500 1.2305 0.11795 0.11445 -0.0424 0.0074 1.0000