XFOIL Version 6.94 Calculated polar for: GOE 682 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 0.4765 0.00847 0.00222 -0.0981 0.6811 0.3569 0.500 0.5169 0.00708 0.00238 -0.0945 0.6609 0.8919 1.000 0.6162 0.00719 0.00233 -0.1037 0.6360 1.0000 1.500 0.6652 0.00740 0.00238 -0.1020 0.6148 1.0000 2.000 0.7144 0.00761 0.00248 -0.1004 0.5927 1.0000 2.500 0.7635 0.00785 0.00260 -0.0988 0.5680 1.0000 3.000 0.8123 0.00812 0.00276 -0.0972 0.5419 1.0000 3.500 0.8610 0.00843 0.00297 -0.0956 0.5127 1.0000 4.000 0.9057 0.00892 0.00321 -0.0933 0.4667 1.0000 4.500 0.9492 0.00955 0.00357 -0.0909 0.4151 1.0000 5.000 0.9730 0.01146 0.00437 -0.0855 0.2395 1.0000 5.500 1.0122 0.01252 0.00506 -0.0827 0.1827 1.0000 6.000 1.0425 0.01417 0.00614 -0.0786 0.0846 1.0000 6.500 1.0689 0.01602 0.00759 -0.0738 0.0040 1.0000 7.000 1.1102 0.01673 0.00834 -0.0713 0.0037 1.0000 7.500 1.1469 0.01751 0.00920 -0.0679 0.0036 1.0000 8.000 1.1804 0.01845 0.01024 -0.0642 0.0036 1.0000 8.500 1.2118 0.01954 0.01153 -0.0603 0.0037 1.0000 9.000 1.2407 0.02082 0.01299 -0.0563 0.0038 1.0000 9.500 1.2653 0.02243 0.01480 -0.0521 0.0039 1.0000 10.000 1.2825 0.02464 0.01724 -0.0474 0.0041 1.0000 10.500 1.2873 0.02793 0.02079 -0.0423 0.0043 1.0000 11.000 1.2934 0.03143 0.02451 -0.0384 0.0045 1.0000 11.500 1.2999 0.03528 0.02859 -0.0354 0.0046 1.0000 12.000 1.3045 0.03967 0.03317 -0.0333 0.0048 1.0000 12.500 1.3080 0.04457 0.03828 -0.0317 0.0054 1.0000 13.000 1.3072 0.05027 0.04422 -0.0307 0.0058 1.0000 13.500 1.3036 0.05666 0.05085 -0.0303 0.0061 1.0000 14.000 1.2969 0.06370 0.05817 -0.0297 0.0071 1.0000 14.500 1.2926 0.07071 0.06544 -0.0294 0.0078 1.0000 15.000 1.2987 0.07644 0.07134 -0.0279 0.0092 1.0000 15.500 1.3067 0.08339 0.07886 -0.0255 0.0132 1.0000