GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Reynolds number: 500,000 Max Cl/Cd: 99.1 at α=7.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe207-il-500000.txt Download as CSV file: xf-goe207-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 207 (AVIATIK V8) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2290 0.09368 0.09129 -0.0207 0.8299 0.0203
-7.250 -0.2198 0.09015 0.08766 -0.0195 0.8044 0.0204
-7.000 -0.2089 0.08715 0.08455 -0.0203 0.7795 0.0206
-6.750 -0.1964 0.08435 0.08164 -0.0220 0.7563 0.0208
-6.500 -0.1818 0.08159 0.07879 -0.0245 0.7346 0.0210
-6.250 -0.1654 0.07878 0.07588 -0.0275 0.7137 0.0214
-6.000 -0.1473 0.07587 0.07287 -0.0309 0.6936 0.0218
-5.750 -0.1273 0.07289 0.06978 -0.0345 0.6737 0.0223
-5.500 -0.1052 0.06985 0.06663 -0.0385 0.6546 0.0229
-5.250 -0.0728 0.06676 0.06339 -0.0454 0.6361 0.0236
-5.000 -0.0348 0.06370 0.06013 -0.0537 0.6184 0.0237
-4.750 -0.0026 0.06034 0.05660 -0.0591 0.6021 0.0238
-4.250 0.0310 0.05431 0.05044 -0.0599 0.5733 0.0241
-4.000 0.0553 0.05179 0.04783 -0.0620 0.5614 0.0244
-3.750 0.0823 0.04939 0.04531 -0.0645 0.5508 0.0248
-3.500 0.1113 0.04701 0.04282 -0.0672 0.5408 0.0253
-3.250 0.1420 0.04463 0.04033 -0.0700 0.5316 0.0261
-3.000 0.1785 0.04240 0.03791 -0.0732 0.5223 0.0274
-2.750 0.2252 0.04049 0.03570 -0.0773 0.5137 0.0277
-2.500 0.2510 0.03747 0.03258 -0.0790 0.5052 0.0280
-2.250 0.2759 0.03538 0.03044 -0.0801 0.4970 0.0283
-2.000 0.3040 0.03362 0.02857 -0.0814 0.4887 0.0286
-1.750 0.3342 0.03195 0.02680 -0.0828 0.4805 0.0292
-1.500 0.3653 0.03037 0.02508 -0.0841 0.4720 0.0300
-1.250 0.3993 0.02885 0.02340 -0.0853 0.4638 0.0317
-1.000 0.4379 0.02696 0.02117 -0.0866 0.4555 0.0327
-0.750 0.4654 0.02544 0.01960 -0.0876 0.4465 0.0332
-0.500 0.4943 0.02431 0.01835 -0.0884 0.4372 0.0340
-0.250 0.5249 0.02322 0.01714 -0.0892 0.4273 0.0353
0.000 0.5616 0.02228 0.01583 -0.0892 0.4182 0.0382
0.250 0.5898 0.02089 0.01437 -0.0903 0.4077 0.0389
0.500 0.6188 0.02005 0.01344 -0.0909 0.3978 0.0399
0.750 0.6486 0.01934 0.01257 -0.0913 0.3882 0.0417
1.000 0.6813 0.01842 0.01137 -0.0914 0.3793 0.0453
1.250 0.7098 0.01773 0.01061 -0.0919 0.3705 0.0466
1.500 0.7392 0.01730 0.01006 -0.0922 0.3619 0.0504
1.750 0.7692 0.01660 0.00917 -0.0924 0.3544 0.0543
2.000 0.7979 0.01621 0.00871 -0.0927 0.3468 0.0576
2.250 0.8273 0.01582 0.00813 -0.0927 0.3396 0.0635
2.500 0.8558 0.01549 0.00777 -0.0930 0.3325 0.0681
2.750 0.8845 0.01521 0.00731 -0.0930 0.3256 0.0742
3.000 0.9133 0.01482 0.00692 -0.0931 0.3198 0.0774
3.250 0.9419 0.01454 0.00654 -0.0931 0.3143 0.0784
3.500 0.9723 0.01334 0.00503 -0.0925 0.3096 0.0435
3.750 1.0009 0.01324 0.00494 -0.0926 0.3054 0.0448
4.000 1.0292 0.01317 0.00489 -0.0928 0.3009 0.0472
4.250 1.0570 0.01326 0.00495 -0.0929 0.2966 0.0492
4.500 1.0849 0.01337 0.00504 -0.0930 0.2927 0.0523
4.750 1.1131 0.01340 0.00513 -0.0931 0.2889 0.0605
5.000 1.1413 0.01342 0.00527 -0.0933 0.2851 0.1335
5.250 1.1620 0.01228 0.00551 -0.0921 0.2807 1.0000
5.500 1.1895 0.01246 0.00566 -0.0921 0.2761 1.0000
5.750 1.2169 0.01263 0.00580 -0.0922 0.2716 1.0000
6.000 1.2439 0.01286 0.00599 -0.0921 0.2671 1.0000
6.250 1.2705 0.01313 0.00622 -0.0921 0.2626 1.0000
6.500 1.2979 0.01327 0.00638 -0.0921 0.2583 1.0000
6.750 1.3246 0.01349 0.00657 -0.0921 0.2527 1.0000
7.000 1.3508 0.01376 0.00681 -0.0921 0.2474 1.0000
7.250 1.3778 0.01392 0.00700 -0.0921 0.2420 1.0000
7.500 1.4036 0.01422 0.00726 -0.0920 0.2352 1.0000
7.750 1.4300 0.01443 0.00747 -0.0919 0.2268 1.0000
8.000 1.4553 0.01477 0.00775 -0.0918 0.2140 1.0000
8.250 1.4785 0.01536 0.00816 -0.0915 0.1861 1.0000
8.500 1.4852 0.01798 0.01013 -0.0894 0.0927 1.0000
9.000 1.5140 0.02084 0.01279 -0.0865 0.0303 1.0000
9.250 1.5318 0.02175 0.01379 -0.0854 0.0278 1.0000
9.500 1.5500 0.02253 0.01466 -0.0843 0.0267 1.0000
9.750 1.5665 0.02340 0.01563 -0.0830 0.0256 1.0000
10.000 1.5800 0.02442 0.01675 -0.0814 0.0246 1.0000
10.250 1.5885 0.02568 0.01811 -0.0793 0.0237 1.0000
10.500 1.5871 0.02733 0.01987 -0.0759 0.0229 1.0000
10.750 1.5800 0.02969 0.02239 -0.0729 0.0224 1.0000
11.000 1.5847 0.03149 0.02429 -0.0717 0.0220 1.0000
11.250 1.5858 0.03391 0.02684 -0.0709 0.0216 1.0000
11.500 1.5842 0.03692 0.02999 -0.0706 0.0213 1.0000
11.750 1.5801 0.04056 0.03376 -0.0709 0.0209 1.0000
12.000 1.5731 0.04478 0.03812 -0.0716 0.0206 1.0000
12.250 1.5634 0.04959 0.04307 -0.0726 0.0204 1.0000
12.500 1.5508 0.05486 0.04848 -0.0737 0.0202 1.0000
12.750 1.5360 0.06048 0.05424 -0.0749 0.0201 1.0000
13.000 1.5195 0.06628 0.06017 -0.0761 0.0200 1.0000
13.250 1.5020 0.07221 0.06623 -0.0772 0.0199 1.0000
13.500 1.4843 0.07819 0.07233 -0.0783 0.0198 1.0000
13.750 1.4683 0.08412 0.07837 -0.0796 0.0197 1.0000
14.000 1.4538 0.08989 0.08424 -0.0809 0.0196 1.0000
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Polar data table (+)
Polar graphs
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