GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 1,000,000 Max Cl/Cd: 111.64 at α=4° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe10k-il-1000000.txt Download as CSV file: xf-goe10k-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 10K 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5396 0.08743 0.08586 -0.0087 1.0000 0.0037
-7.500 -0.5450 0.08481 0.08326 -0.0085 1.0000 0.0036
-7.250 -0.5451 0.08125 0.07973 -0.0103 1.0000 0.0038
-7.000 -0.5417 0.07758 0.07606 -0.0127 1.0000 0.0037
-6.750 -0.5366 0.07362 0.07210 -0.0154 1.0000 0.0039
-6.500 -0.5295 0.06957 0.06803 -0.0179 1.0000 0.0039
-6.250 -0.5207 0.06539 0.06382 -0.0203 1.0000 0.0041
-6.000 -0.5093 0.06084 0.05922 -0.0226 1.0000 0.0043
-5.750 -0.4959 0.05641 0.05472 -0.0244 1.0000 0.0044
-5.500 -0.4809 0.05190 0.05011 -0.0256 1.0000 0.0045
-5.250 -0.4648 0.04740 0.04549 -0.0261 1.0000 0.0045
-5.000 -0.4349 0.04211 0.04002 -0.0293 0.9988 0.0046
-4.750 -0.4066 0.03621 0.03390 -0.0330 0.9971 0.0048
-4.500 -0.3767 0.03249 0.02998 -0.0354 0.9957 0.0051
-4.250 -0.3466 0.02841 0.02564 -0.0368 0.9942 0.0056
-4.000 -0.3144 0.02373 0.02059 -0.0364 0.9922 0.0065
-3.750 -0.2868 0.01890 0.01527 -0.0367 0.9904 0.0069
-3.500 -0.2575 0.01705 0.01319 -0.0375 0.9891 0.0078
-3.250 -0.2271 0.01491 0.01069 -0.0378 0.9881 0.0109
-3.000 -0.1964 0.01312 0.00857 -0.0382 0.9873 0.0139
-2.750 -0.1685 0.01215 0.00749 -0.0381 0.9853 0.0158
-2.500 -0.1426 0.00887 0.00367 -0.0357 0.9839 0.0062
-2.250 -0.1139 0.00804 0.00273 -0.0358 0.9821 0.0075
-2.000 -0.0831 0.00758 0.00222 -0.0365 0.9806 0.0107
-1.750 -0.0508 0.00710 0.00178 -0.0374 0.9793 0.0379
-1.500 -0.0166 0.00684 0.00163 -0.0390 0.9782 0.0801
-1.250 0.0114 0.00641 0.00150 -0.0393 0.9746 0.1749
-1.000 0.0382 0.00527 0.00139 -0.0397 0.9708 0.5172
-0.750 0.0638 0.00419 0.00132 -0.0393 0.9679 0.8310
-0.500 0.1479 0.00408 0.00151 -0.0523 0.9777 0.9975
-0.250 0.1811 0.00401 0.00141 -0.0536 0.9732 1.0000
0.000 0.2169 0.00390 0.00129 -0.0554 0.9694 1.0000
0.250 0.2528 0.00373 0.00112 -0.0571 0.9601 1.0000
0.500 0.2983 0.00356 0.00095 -0.0611 0.9443 1.0000
0.750 0.3501 0.00346 0.00081 -0.0665 0.9173 1.0000
1.000 0.3829 0.00354 0.00076 -0.0675 0.8721 1.0000
1.250 0.4011 0.00385 0.00075 -0.0650 0.7831 1.0000
1.500 0.4024 0.00515 0.00090 -0.0593 0.4879 1.0000
1.750 0.4124 0.00650 0.00117 -0.0560 0.1911 1.0000
2.000 0.4318 0.00715 0.00139 -0.0544 0.0754 1.0000
2.250 0.4536 0.00759 0.00162 -0.0532 0.0170 1.0000
2.500 0.4760 0.00805 0.00212 -0.0519 0.0090 1.0000
2.750 0.4974 0.00866 0.00280 -0.0503 0.0069 1.0000
3.000 0.5197 0.00918 0.00339 -0.0490 0.0062 1.0000
3.250 0.5412 0.00982 0.00407 -0.0475 0.0052 1.0000
3.750 2.1728 0.59188 0.59038 -0.5010 0.0069 1.0000
4.000 0.5783 0.00518 0.00044 -0.0361 0.0098 1.0000
4.250 0.6023 0.00544 0.00104 -0.0344 0.0082 1.0000
4.500 0.6216 0.00662 0.00239 -0.0329 0.0075 1.0000
4.750 0.6338 0.00965 0.00575 -0.0303 0.0070 1.0000
5.000 0.6548 0.01183 0.00838 -0.0272 0.0062 1.0000
5.250 0.6704 0.01461 0.01144 -0.0248 0.0056 1.0000
5.500 0.6847 0.01702 0.01403 -0.0233 0.0052 1.0000
5.750 0.6898 0.02160 0.01887 -0.0211 0.0049 1.0000
6.000 0.7052 0.02607 0.02363 -0.0183 0.0048 1.0000
6.250 0.7168 0.03119 0.02898 -0.0160 0.0042 1.0000
6.500 0.7246 0.03602 0.03398 -0.0145 0.0039 1.0000
6.750 0.7303 0.04083 0.03894 -0.0133 0.0037 1.0000
7.000 0.7337 0.04553 0.04377 -0.0125 0.0036 1.0000
7.250 0.7342 0.05030 0.04864 -0.0119 0.0035 1.0000
7.500 0.7317 0.05522 0.05365 -0.0116 0.0035 1.0000
7.750 0.7234 0.05965 0.05813 -0.0113 0.0033 1.0000
8.000 0.7118 0.06414 0.06268 -0.0109 0.0034 1.0000
8.250 0.6895 0.07025 0.06883 -0.0141 0.0036 1.0000
8.500 0.6745 0.07621 0.07477 -0.0185 0.0037 1.0000
8.750 0.6659 0.08171 0.08026 -0.0212 0.0037 1.0000
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Polar data table (+)
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