XFOIL Version 6.96 Calculated polar for: GOE 443 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6438 0.08678 0.08460 0.0054 1.0000 0.0071 -8.000 -0.6468 0.08263 0.08048 0.0025 1.0000 0.0072 -7.750 -0.6512 0.07804 0.07590 -0.0015 1.0000 0.0072 -7.500 -0.6492 0.07338 0.07121 -0.0050 1.0000 0.0073 -7.250 -0.6454 0.06859 0.06636 -0.0079 1.0000 0.0074 -7.000 -0.6391 0.06412 0.06182 -0.0100 1.0000 0.0076 -6.750 -0.6309 0.05968 0.05727 -0.0115 1.0000 0.0078 -6.500 -0.6205 0.05534 0.05280 -0.0125 1.0000 0.0081 -6.250 -0.6079 0.05121 0.04851 -0.0130 1.0000 0.0085 -6.000 -0.5934 0.04714 0.04427 -0.0131 1.0000 0.0089 -5.750 -0.5762 0.04332 0.04023 -0.0126 1.0000 0.0096 -5.500 -0.5517 0.04060 0.03726 -0.0114 1.0000 0.0106 -5.250 -0.5322 0.03764 0.03404 -0.0102 1.0000 0.0108 -5.000 -0.5139 0.03446 0.03061 -0.0090 1.0000 0.0108 -4.750 -0.4953 0.03134 0.02720 -0.0076 1.0000 0.0109 -4.500 -0.4762 0.02838 0.02395 -0.0061 1.0000 0.0109 -4.250 -0.4565 0.02551 0.02078 -0.0046 1.0000 0.0109 -4.000 -0.4398 0.01922 0.01402 -0.0020 1.0000 0.0059 -3.750 -0.4176 0.01675 0.01122 -0.0004 1.0000 0.0055 -3.500 -0.3943 0.01463 0.00880 0.0011 1.0000 0.0050 -3.250 -0.3709 0.01287 0.00681 0.0026 1.0000 0.0047 -3.000 -0.3484 0.01139 0.00515 0.0041 1.0000 0.0043 -2.750 -0.3263 0.01033 0.00377 0.0056 1.0000 0.0040 -2.500 -0.3040 0.00948 0.00280 0.0069 1.0000 0.0038 -2.250 -0.2807 0.00895 0.00215 0.0080 1.0000 0.0036 -2.000 -0.2568 0.00862 0.00171 0.0089 1.0000 0.0035 -1.750 -0.2297 0.00839 0.00139 0.0090 0.9991 0.0034 -1.500 -0.1936 0.00821 0.00114 0.0071 0.9943 0.0035 -1.250 -0.1591 0.00807 0.00097 0.0055 0.9889 0.0038 -1.000 -0.1261 0.00797 0.00087 0.0042 0.9835 0.0050 -0.750 -0.1035 0.00585 0.00074 0.0040 0.9767 0.5410 -0.500 -0.0851 0.00461 0.00091 0.0064 0.9695 0.8891 -0.250 -0.0454 0.00459 0.00097 0.0039 0.9657 0.9354 0.000 -0.0002 0.00461 0.00101 0.0000 0.9602 0.9600 0.250 0.0454 0.00459 0.00098 -0.0039 0.9357 0.9657 0.500 0.0852 0.00461 0.00091 -0.0065 0.8896 0.9695 0.750 0.1041 0.00580 0.00074 -0.0041 0.5550 0.9765 1.000 0.1262 0.00797 0.00087 -0.0042 0.0050 0.9835 1.250 0.1591 0.00807 0.00097 -0.0055 0.0039 0.9889 1.500 0.1938 0.00821 0.00114 -0.0071 0.0035 0.9943 1.750 0.2302 0.00839 0.00139 -0.0091 0.0034 0.9991 2.000 0.2573 0.00862 0.00171 -0.0090 0.0035 1.0000 2.250 0.2813 0.00896 0.00217 -0.0081 0.0036 1.0000 2.500 0.3046 0.00951 0.00283 -0.0070 0.0038 1.0000 2.750 0.3269 0.01035 0.00379 -0.0057 0.0040 1.0000 3.000 0.3490 0.01141 0.00517 -0.0043 0.0043 1.0000 3.250 0.3715 0.01284 0.00679 -0.0027 0.0046 1.0000 3.500 0.3949 0.01465 0.00883 -0.0012 0.0050 1.0000 3.750 0.4182 0.01678 0.01125 0.0003 0.0055 1.0000 4.000 0.4402 0.01928 0.01408 0.0019 0.0060 1.0000 4.250 0.4569 0.02554 0.02081 0.0045 0.0109 1.0000 4.750 0.4721 0.01614 0.01229 0.0090 0.0109 1.0000 5.000 0.4894 0.01920 0.01562 0.0102 0.0109 1.0000 5.250 0.5061 0.02251 0.01919 0.0113 0.0108 1.0000 5.500 0.5226 0.02595 0.02288 0.0122 0.0107 1.0000 5.750 0.5426 0.02888 0.02604 0.0131 0.0102 1.0000 6.000 0.5598 0.03270 0.03008 0.0136 0.0092 1.0000 6.250 0.5712 0.03729 0.03484 0.0134 0.0087 1.0000 6.500 0.5794 0.04219 0.03989 0.0128 0.0084 1.0000 6.750 0.5848 0.04723 0.04505 0.0117 0.0082 1.0000 7.000 0.5871 0.05229 0.05020 0.0101 0.0080 1.0000 7.250 0.5857 0.05751 0.05548 0.0077 0.0079 1.0000 7.500 0.5780 0.06257 0.06058 0.0045 0.0079 1.0000 7.750 0.5659 0.06723 0.06521 0.0009 0.0080 1.0000 8.000 0.5592 0.07197 0.06993 -0.0016 0.0080 1.0000 8.250 0.5552 0.07654 0.07448 -0.0036 0.0079 1.0000 8.500 0.5519 0.08123 0.07916 -0.0054 0.0078 1.0000