GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 50,000 Max Cl/Cd: 22.69 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe10k-il-50000-n5.txt Download as CSV file: xf-goe10k-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.5435 0.09742 0.09065 -0.0114 1.0000 0.0908
-7.500 -0.5453 0.09475 0.08807 -0.0146 1.0000 0.0942
-7.250 -0.5486 0.09271 0.08609 -0.0225 1.0000 0.0963
-7.000 -0.5387 0.08715 0.08057 -0.0184 1.0000 0.0985
-6.750 -0.5316 0.08326 0.07668 -0.0184 1.0000 0.1010
-6.500 -0.5248 0.07957 0.07301 -0.0199 1.0000 0.1040
-6.250 -0.5187 0.07610 0.06948 -0.0252 1.0000 0.1115
-5.750 -0.4848 0.06543 0.05856 -0.0288 1.0000 0.0618
-5.500 -0.4649 0.06011 0.05298 -0.0312 1.0000 0.0429
-5.250 -0.4507 0.05592 0.04864 -0.0318 1.0000 0.0408
-5.000 -0.4291 0.05116 0.04340 -0.0336 1.0000 0.0367
-4.750 -0.4121 0.04738 0.03941 -0.0335 1.0000 0.0362
-4.500 -0.3932 0.04371 0.03542 -0.0335 1.0000 0.0359
-4.250 -0.3729 0.04016 0.03150 -0.0332 1.0000 0.0354
-4.000 -0.3514 0.03681 0.02771 -0.0326 1.0000 0.0348
-3.750 -0.3286 0.03359 0.02396 -0.0319 1.0000 0.0344
-3.500 -0.3045 0.03067 0.02046 -0.0309 1.0000 0.0349
-3.250 -0.2788 0.02805 0.01712 -0.0297 1.0000 0.0367
-3.000 -0.2550 0.02587 0.01470 -0.0288 1.0000 0.0399
-2.750 -0.2289 0.02389 0.01223 -0.0276 1.0000 0.0437
-2.500 -0.2026 0.02211 0.01021 -0.0267 1.0000 0.0517
-2.250 -0.1761 0.02066 0.00837 -0.0257 1.0000 0.0654
-2.000 -0.1504 0.01913 0.00686 -0.0251 1.0000 0.1017
-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 0.5068 0.02234 0.00960 -0.0324 0.0615 1.0000
3.750 0.5340 0.02415 0.01160 -0.0318 0.0466 1.0000
4.000 0.5645 0.02614 0.01393 -0.0312 0.0395 1.0000
4.250 0.5943 0.02848 0.01662 -0.0305 0.0346 1.0000
4.500 0.6226 0.03120 0.01978 -0.0295 0.0330 1.0000
4.750 0.6473 0.03397 0.02296 -0.0283 0.0311 1.0000
5.000 0.6688 0.03714 0.02650 -0.0270 0.0293 1.0000
5.250 0.6895 0.04045 0.03044 -0.0252 0.0289 1.0000
5.500 0.7075 0.04403 0.03449 -0.0235 0.0292 1.0000
5.750 0.7235 0.04784 0.03867 -0.0220 0.0296 1.0000
6.000 0.7393 0.05166 0.04326 -0.0198 0.0313 1.0000
6.250 0.7513 0.05608 0.04816 -0.0183 0.0330 1.0000
6.500 0.7611 0.06043 0.05284 -0.0172 0.0345 1.0000
6.750 0.7693 0.06473 0.05739 -0.0165 0.0358 1.0000
7.000 0.7764 0.06899 0.06180 -0.0158 0.0371 1.0000
7.500 0.7825 0.07861 0.07185 -0.0162 0.0435 1.0000
8.000 0.7117 0.07869 0.07242 -0.0103 0.0458 1.0000
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