GOE 509 AIRFOIL (goe509-il) Xfoil prediction polar at RE=100,000 Ncrit=9
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Airfoil: GOE 509 AIRFOIL (goe509-il) Reynolds number: 100,000 Max Cl/Cd: 51.58 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe509-il-100000.txt Download as CSV file: xf-goe509-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 509 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.3672 0.09597 0.09187 -0.0213 1.0000 0.0907
-7.250 -0.3897 0.09523 0.09124 -0.0178 1.0000 0.0911
-7.000 -0.4079 0.09410 0.09017 -0.0151 1.0000 0.0921
-6.750 -0.4226 0.09254 0.08868 -0.0134 1.0000 0.0938
-6.500 -0.4278 0.09026 0.08639 -0.0158 0.9977 0.0968
-6.250 -0.4073 0.08459 0.08064 -0.0247 0.9885 0.1014
-6.000 -0.3803 0.08103 0.07707 -0.0260 0.9834 0.1071
-5.750 -0.3566 0.07639 0.07216 -0.0376 0.9718 0.1159
-5.500 -0.3385 0.07263 0.06848 -0.0360 0.9646 0.1190
-5.250 -0.2426 0.05561 0.05170 -0.0457 0.9333 0.1368
-5.000 -0.2305 0.05183 0.04776 -0.0489 0.9249 0.1488
-4.750 -0.2671 0.06217 0.05764 -0.0457 0.9380 0.1490
-4.500 -0.2387 0.05915 0.05448 -0.0484 0.9306 0.1642
-4.000 -0.1897 0.05395 0.04902 -0.0509 0.9151 0.1949
-3.750 -0.1738 0.05192 0.04692 -0.0499 0.9050 0.2106
-3.500 -0.1404 0.04882 0.04383 -0.0513 0.9001 0.2302
-3.250 -0.1159 0.03717 0.03042 -0.0521 0.8906 0.1104
-3.000 -0.0795 0.03282 0.02503 -0.0526 0.8860 0.1015
-2.750 -0.0569 0.03108 0.02302 -0.0513 0.8761 0.1028
-2.500 -0.0118 0.02895 0.02053 -0.0537 0.8715 0.1055
-2.250 0.0151 0.02779 0.01911 -0.0529 0.8610 0.1099
-2.000 0.0671 0.02611 0.01698 -0.0561 0.8567 0.1146
-1.750 0.0978 0.02477 0.01570 -0.0563 0.8461 0.1221
-1.500 0.1533 0.02311 0.01393 -0.0603 0.8422 0.1328
-1.250 0.1881 0.02212 0.01300 -0.0610 0.8321 0.1475
-1.000 0.2441 0.02055 0.01174 -0.0654 0.8273 0.1842
-0.750 0.2821 0.01963 0.01139 -0.0663 0.8164 0.3286
-0.500 0.5274 0.01632 0.00962 -0.1086 0.8181 1.0000
-0.250 0.5643 0.01602 0.00906 -0.1096 0.8017 1.0000
0.000 0.5969 0.01583 0.00864 -0.1098 0.7824 1.0000
0.250 0.6276 0.01573 0.00830 -0.1096 0.7621 1.0000
0.500 0.6502 0.01582 0.00820 -0.1080 0.7390 1.0000
0.750 0.6735 0.01591 0.00810 -0.1067 0.7171 1.0000
1.000 0.6978 0.01600 0.00798 -0.1055 0.6975 1.0000
1.250 0.7202 0.01615 0.00795 -0.1041 0.6788 1.0000
1.500 0.7415 0.01634 0.00801 -0.1026 0.6615 1.0000
1.750 0.7635 0.01654 0.00805 -0.1012 0.6458 1.0000
2.000 0.7865 0.01672 0.00807 -0.1001 0.6321 1.0000
2.250 0.8070 0.01695 0.00820 -0.0985 0.6178 1.0000
2.500 0.8273 0.01719 0.00837 -0.0970 0.6046 1.0000
2.750 0.8481 0.01743 0.00853 -0.0955 0.5922 1.0000
3.000 0.8697 0.01765 0.00864 -0.0941 0.5802 1.0000
3.250 0.8928 0.01785 0.00870 -0.0930 0.5692 1.0000
3.500 0.9117 0.01814 0.00897 -0.0912 0.5573 1.0000
3.750 0.9328 0.01844 0.00925 -0.0899 0.5475 1.0000
4.000 0.9563 0.01874 0.00945 -0.0890 0.5389 1.0000
4.250 0.9758 0.01912 0.00987 -0.0875 0.5299 1.0000
4.500 1.0010 0.01943 0.01007 -0.0869 0.5221 1.0000
4.750 1.0189 0.01986 0.01059 -0.0851 0.5129 1.0000
5.000 1.0430 0.02022 0.01087 -0.0843 0.5048 1.0000
5.250 1.0619 0.02064 0.01134 -0.0827 0.4954 1.0000
5.500 1.0834 0.02107 0.01178 -0.0815 0.4864 1.0000
5.750 1.1059 0.02148 0.01215 -0.0805 0.4769 1.0000
6.000 1.1236 0.02200 0.01273 -0.0786 0.4668 1.0000
6.250 1.1473 0.02248 0.01315 -0.0779 0.4572 1.0000
6.500 1.1654 0.02301 0.01375 -0.0761 0.4465 1.0000
6.750 1.1833 0.02362 0.01442 -0.0743 0.4358 1.0000
7.000 1.2065 0.02420 0.01494 -0.0735 0.4251 1.0000
7.250 1.2232 0.02474 0.01553 -0.0715 0.4128 1.0000
7.500 1.2370 0.02533 0.01624 -0.0690 0.4001 1.0000
7.750 1.2532 0.02586 0.01679 -0.0669 0.3877 1.0000
8.000 1.2719 0.02625 0.01712 -0.0652 0.3757 1.0000
8.250 1.2852 0.02665 0.01761 -0.0626 0.3639 1.0000
8.500 1.2961 0.02715 0.01825 -0.0597 0.3527 1.0000
8.750 1.3127 0.02752 0.01862 -0.0577 0.3429 1.0000
9.000 1.3254 0.02786 0.01904 -0.0551 0.3330 1.0000
9.250 1.3357 0.02836 0.01966 -0.0521 0.3238 1.0000
9.500 1.3536 0.02858 0.01986 -0.0504 0.3157 1.0000
9.750 1.3580 0.02908 0.02055 -0.0465 0.3063 1.0000
10.000 1.3689 0.02933 0.02081 -0.0436 0.2979 1.0000
10.250 1.3756 0.02976 0.02136 -0.0402 0.2894 1.0000
10.500 1.3829 0.03022 0.02190 -0.0369 0.2819 1.0000
10.750 1.3881 0.03073 0.02251 -0.0333 0.2747 1.0000
11.000 1.3939 0.03126 0.02312 -0.0299 0.2678 1.0000
11.250 1.3967 0.03187 0.02385 -0.0262 0.2603 1.0000
11.500 1.3971 0.03252 0.02458 -0.0224 0.2518 1.0000
11.750 1.3983 0.03321 0.02531 -0.0190 0.2432 1.0000
12.000 1.3972 0.03424 0.02652 -0.0156 0.2348 1.0000
12.250 1.3949 0.03529 0.02760 -0.0123 0.2251 1.0000
12.500 1.3888 0.03670 0.02912 -0.0090 0.2143 1.0000
12.750 1.3818 0.03847 0.03105 -0.0062 0.2033 1.0000
13.000 1.3738 0.04055 0.03323 -0.0038 0.1917 1.0000
13.250 1.3631 0.04310 0.03587 -0.0018 0.1788 1.0000
13.500 1.3489 0.04630 0.03916 -0.0003 0.1639 1.0000
13.750 1.3318 0.05019 0.04316 0.0007 0.1466 1.0000
14.000 1.3136 0.05467 0.04771 0.0010 0.1264 1.0000
14.250 1.2933 0.05976 0.05276 0.0008 0.1066 1.0000
14.500 1.2705 0.06547 0.05838 0.0000 0.0970 1.0000
14.750 1.2490 0.07136 0.06421 -0.0011 0.0911 1.0000
15.000 1.2302 0.07708 0.06991 -0.0022 0.0859 1.0000
15.250 1.2135 0.08257 0.07537 -0.0034 0.0823 1.0000
15.500 1.2029 0.08723 0.08007 -0.0041 0.0781 1.0000
15.750 1.1932 0.09176 0.08465 -0.0049 0.0739 1.0000
16.000 1.1894 0.09503 0.08783 -0.0049 0.0705 1.0000
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Polar data table (+)
Polar graphs
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