GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 50,000 Max Cl/Cd: 20.6 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe10k-il-50000.txt Download as CSV file: xf-goe10k-il-50000.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 = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5491 0.10529 0.09845 -0.0053 1.0000 0.1573
-8.000 -0.5542 0.10329 0.09655 -0.0065 1.0000 0.1639
-7.750 -0.5702 0.10272 0.09614 -0.0108 1.0000 0.1659
-7.500 -0.5532 0.09679 0.09018 -0.0068 1.0000 0.1778
-7.250 -0.5501 0.09296 0.08642 -0.0069 1.0000 0.1854
-7.000 -0.5566 0.09132 0.08487 -0.0113 1.0000 0.1949
-6.750 -0.5478 0.08685 0.08044 -0.0090 1.0000 0.2085
-6.500 -0.5435 0.08329 0.07693 -0.0088 1.0000 0.2230
-6.250 -0.5435 0.08062 0.07430 -0.0111 1.0000 0.2399
-6.000 -0.5338 0.07621 0.06997 -0.0062 1.0000 0.2642
-5.750 -0.5298 0.07290 0.06673 -0.0046 1.0000 0.2908
-5.500 -0.5248 0.06978 0.06367 -0.0016 1.0000 0.3242
-5.250 -0.5214 0.06668 0.06065 0.0027 1.0000 0.3667
-4.500 -0.3662 0.05049 0.04405 0.0275 1.0000 0.7976
-4.250 -0.3902 0.04951 0.04326 0.0316 1.0000 0.7756
-4.000 -0.4191 0.04844 0.04240 0.0351 1.0000 0.7357
-3.750 -0.3525 0.03868 0.03018 -0.0322 1.0000 0.1487
-3.500 -0.3219 0.03487 0.02560 -0.0322 1.0000 0.1215
-3.250 -0.2935 0.03184 0.02179 -0.0313 1.0000 0.1099
-3.000 -0.2680 0.02874 0.01835 -0.0304 1.0000 0.1059
-2.750 -0.2408 0.02628 0.01534 -0.0292 1.0000 0.1069
-2.500 -0.2133 0.02399 0.01259 -0.0281 1.0000 0.1132
-2.250 -0.1853 0.02194 0.01035 -0.0271 1.0000 0.1312
-2.000 -0.1547 0.01978 0.00807 -0.0263 1.0000 0.1806
-1.750 -0.1009 0.01462 0.00535 -0.0290 1.0000 1.0000
-1.500 -0.0780 0.01461 0.00464 -0.0278 1.0000 1.0000
-1.250 -0.0554 0.01461 0.00418 -0.0268 1.0000 1.0000
-1.000 -0.0329 0.01464 0.00383 -0.0258 1.0000 1.0000
-0.750 -0.0104 0.01468 0.00356 -0.0248 1.0000 1.0000
-0.500 0.0121 0.01473 0.00332 -0.0238 1.0000 1.0000
-0.250 0.0344 0.01480 0.00320 -0.0229 1.0000 1.0000
0.000 0.0568 0.01489 0.00314 -0.0220 1.0000 1.0000
0.250 0.0792 0.01499 0.00311 -0.0211 1.0000 1.0000
0.500 0.1015 0.01511 0.00316 -0.0203 1.0000 1.0000
0.750 0.1237 0.01524 0.00325 -0.0194 1.0000 1.0000
1.000 0.1459 0.01539 0.00340 -0.0186 1.0000 1.0000
1.250 0.1680 0.01555 0.00361 -0.0177 1.0000 1.0000
1.500 0.1899 0.01573 0.00387 -0.0169 1.0000 1.0000
1.750 0.2119 0.01593 0.00418 -0.0161 1.0000 1.0000
2.000 0.2336 0.01616 0.00459 -0.0152 1.0000 1.0000
2.250 0.2553 0.01640 0.00503 -0.0144 1.0000 1.0000
2.500 0.2768 0.01667 0.00554 -0.0136 1.0000 1.0000
2.750 0.2982 0.01697 0.00616 -0.0128 1.0000 1.0000
3.000 0.3193 0.01730 0.00683 -0.0119 1.0000 1.0000
3.250 0.3402 0.01767 0.00760 -0.0111 1.0000 1.0000
3.500 0.3611 0.01808 0.00849 -0.0103 1.0000 1.0000
3.750 0.3817 0.01853 0.00962 -0.0094 1.0000 1.0000
4.000 0.5878 0.02881 0.01711 -0.0311 0.0935 1.0000
4.250 0.6166 0.03159 0.02038 -0.0299 0.0899 1.0000
4.500 0.6440 0.03490 0.02424 -0.0284 0.0943 1.0000
4.750 0.6696 0.03821 0.02811 -0.0266 0.1007 1.0000
5.000 0.6927 0.04217 0.03266 -0.0248 0.1108 1.0000
5.250 0.7170 0.04623 0.03762 -0.0230 0.1349 1.0000
5.500 0.7587 0.05384 0.04708 -0.0289 0.2792 1.0000
5.750 0.6807 0.04889 0.04347 -0.0451 0.4409 1.0000
6.000 0.6367 0.05424 0.04864 -0.0491 0.4748 1.0000
7.000 0.7524 0.08577 0.07964 -0.0588 0.4549 1.0000
7.250 0.7505 0.08842 0.08221 -0.0543 0.4064 1.0000
7.500 0.7636 0.09206 0.08584 -0.0500 0.3610 1.0000
7.750 0.7878 0.09669 0.09051 -0.0449 0.3175 1.0000
8.000 0.7703 0.09926 0.09295 -0.0454 0.2996 1.0000
8.250 0.7875 0.10408 0.09781 -0.0421 0.2716 1.0000
8.500 0.7810 0.10746 0.10110 -0.0424 0.2560 1.0000
8.750 0.7782 0.11117 0.10475 -0.0426 0.2407 1.0000
9.000 0.7805 0.11530 0.10884 -0.0424 0.2262 1.0000
9.250 0.7872 0.11989 0.11341 -0.0417 0.2120 1.0000
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