XFOIL Version 6.96 Calculated polar for: GOE 444 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 ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6344 0.08204 0.08052 -0.0029 1.0000 0.0039 -8.000 -0.6432 0.07711 0.07562 -0.0076 1.0000 0.0039 -7.750 -0.6471 0.07202 0.07050 -0.0109 1.0000 0.0035 -7.500 -0.6461 0.06730 0.06573 -0.0134 1.0000 0.0038 -7.250 -0.6429 0.06271 0.06106 -0.0150 1.0000 0.0041 -7.000 -0.6381 0.05815 0.05640 -0.0159 1.0000 0.0041 -6.750 -0.6313 0.05371 0.05179 -0.0160 1.0000 0.0043 -6.500 -0.6225 0.04950 0.04744 -0.0156 1.0000 0.0045 -6.250 -0.6121 0.04547 0.04325 -0.0147 1.0000 0.0048 -6.000 -0.6001 0.04161 0.03921 -0.0134 1.0000 0.0050 -5.750 -0.5860 0.03801 0.03539 -0.0117 1.0000 0.0054 -5.500 -0.5655 0.03515 0.03232 -0.0095 1.0000 0.0061 -5.250 -0.5480 0.03251 0.02945 -0.0075 1.0000 0.0062 -5.000 -0.5315 0.02977 0.02647 -0.0053 1.0000 0.0062 -4.750 -0.5149 0.02718 0.02362 -0.0030 1.0000 0.0063 -4.500 -0.4979 0.02478 0.02092 -0.0007 1.0000 0.0063 -2.750 -0.3126 0.00946 0.00400 0.0024 0.9891 0.0049 -2.500 -0.2821 0.00821 0.00260 0.0021 0.9862 0.0036 -2.250 -0.2530 0.00762 0.00189 0.0019 0.9807 0.0030 -2.000 -0.2197 0.00727 0.00146 0.0008 0.9770 0.0026 -1.750 -0.1895 0.00705 0.00118 0.0002 0.9697 0.0025 -1.500 -0.1557 0.00688 0.00096 -0.0012 0.9633 0.0024 -1.250 -0.1207 0.00675 0.00081 -0.0029 0.9536 0.0025 -1.000 -0.0852 0.00664 0.00069 -0.0047 0.9389 0.0034 -0.750 -0.0611 0.00519 0.00051 -0.0047 0.9117 0.3985 -0.500 -0.0371 0.00476 0.00041 -0.0041 0.8792 0.5275 -0.250 -0.0169 0.00436 0.00038 -0.0025 0.8466 0.6706 0.000 0.0001 0.00418 0.00038 0.0000 0.7839 0.7825 0.250 0.0177 0.00431 0.00038 0.0024 0.6849 0.8459 0.500 0.0375 0.00472 0.00041 0.0040 0.5384 0.8790 0.750 0.0613 0.00517 0.00050 0.0047 0.4050 0.9112 1.000 0.0852 0.00664 0.00069 0.0047 0.0034 0.9390 1.250 0.1208 0.00675 0.00081 0.0028 0.0026 0.9538 1.500 0.1556 0.00688 0.00096 0.0012 0.0024 0.9633 1.750 0.1896 0.00705 0.00117 -0.0003 0.0024 0.9698 2.000 0.2194 0.00727 0.00146 -0.0007 0.0026 0.9767 2.250 0.2525 0.00761 0.00188 -0.0019 0.0030 0.9805 2.500 0.2817 0.00827 0.00267 -0.0020 0.0037 0.9860 2.750 0.3120 0.00947 0.00401 -0.0023 0.0050 0.9890 3.000 0.3415 0.01070 0.00545 -0.0024 0.0054 0.9916 3.250 0.3677 0.01517 0.01014 -0.0011 0.0106 0.9936 3.500 0.3970 0.01718 0.01238 -0.0011 0.0106 0.9955 3.750 0.4273 0.01892 0.01441 -0.0013 0.0105 0.9974 4.000 0.4610 0.01881 0.01450 -0.0014 0.0085 0.9988 4.250 0.4884 0.02038 0.01627 -0.0016 0.0072 1.0000 4.500 0.4984 0.02482 0.02097 0.0007 0.0063 1.0000 4.750 0.5153 0.02719 0.02363 0.0029 0.0063 1.0000 5.000 0.5318 0.02979 0.02650 0.0052 0.0062 1.0000 5.250 0.5483 0.03251 0.02946 0.0074 0.0062 1.0000 5.500 0.5653 0.03523 0.03240 0.0095 0.0061 1.0000 5.750 0.5855 0.03800 0.03539 0.0116 0.0056 1.0000 6.000 0.6001 0.04161 0.03920 0.0134 0.0050 1.0000 6.250 0.6121 0.04548 0.04326 0.0148 0.0046 1.0000 6.500 0.6223 0.04945 0.04740 0.0157 0.0046 1.0000 6.750 0.6310 0.05368 0.05177 0.0161 0.0044 1.0000 7.000 0.6378 0.05806 0.05627 0.0160 0.0041 1.0000 7.250 0.6427 0.06259 0.06094 0.0151 0.0041 1.0000 7.500 0.6452 0.06742 0.06584 0.0135 0.0039 1.0000 7.750 0.6464 0.07208 0.07056 0.0110 0.0037 1.0000 8.000 0.6427 0.07707 0.07558 0.0077 0.0040 1.0000 8.250 0.6342 0.08192 0.08041 0.0030 0.0041 1.0000