GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 10K AIRFOIL (goe10k-il) Reynolds number: 100,000 Max Cl/Cd: 33.1 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe10k-il-100000-n5.txt Download as CSV file: xf-goe10k-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5364 0.09961 0.09462 -0.0122 1.0000 0.0362
-8.000 -0.5388 0.09687 0.09197 -0.0142 1.0000 0.0365
-7.750 -0.5407 0.09380 0.08897 -0.0172 1.0000 0.0366
-7.500 -0.5371 0.09024 0.08543 -0.0212 1.0000 0.0368
-7.250 -0.5334 0.08573 0.08098 -0.0194 1.0000 0.0374
-7.000 -0.5283 0.08211 0.07738 -0.0186 1.0000 0.0385
-6.750 -0.5225 0.07847 0.07375 -0.0200 1.0000 0.0399
-6.500 -0.5147 0.07469 0.06994 -0.0224 1.0000 0.0413
-6.250 -0.4982 0.07089 0.06600 -0.0293 1.0000 0.0442
-6.000 -0.4846 0.06676 0.06165 -0.0325 1.0000 0.0448
-5.750 -0.4784 0.06192 0.05690 -0.0312 1.0000 0.0457
-5.500 -0.4674 0.05806 0.05299 -0.0311 1.0000 0.0466
-5.250 -0.4538 0.05422 0.04906 -0.0315 1.0000 0.0478
-5.000 -0.4316 0.04852 0.04304 -0.0323 1.0000 0.0261
-4.750 -0.4148 0.04455 0.03887 -0.0324 1.0000 0.0245
-4.500 -0.3956 0.04053 0.03457 -0.0323 1.0000 0.0227
-4.250 -0.3747 0.03648 0.03014 -0.0319 1.0000 0.0212
-4.000 -0.3509 0.03217 0.02524 -0.0308 1.0000 0.0190
-3.750 -0.3272 0.02964 0.02223 -0.0294 1.0000 0.0180
-3.500 -0.3053 0.02664 0.01874 -0.0283 1.0000 0.0176
-3.250 -0.2822 0.02383 0.01542 -0.0271 1.0000 0.0175
-3.000 -0.2580 0.02132 0.01241 -0.0260 1.0000 0.0182
-2.750 -0.2342 0.01928 0.01006 -0.0251 1.0000 0.0211
-2.500 -0.2095 0.01759 0.00810 -0.0239 1.0000 0.0228
-2.250 -0.1857 0.01633 0.00658 -0.0226 1.0000 0.0270
-2.000 -0.1626 0.01526 0.00542 -0.0215 1.0000 0.0347
-1.750 -0.1392 0.01446 0.00458 -0.0204 1.0000 0.0575
-1.500 -0.1151 0.01316 0.00395 -0.0201 1.0000 0.2029
-1.250 -0.0586 0.01076 0.00352 -0.0257 1.0000 1.0000
-1.000 -0.0362 0.01080 0.00327 -0.0246 1.0000 1.0000
-0.750 -0.0139 0.01086 0.00309 -0.0236 1.0000 1.0000
-0.500 0.0085 0.01093 0.00293 -0.0227 1.0000 1.0000
-0.250 0.0307 0.01101 0.00287 -0.0217 1.0000 1.0000
0.000 0.0529 0.01111 0.00286 -0.0208 1.0000 1.0000
0.250 0.0751 0.01122 0.00287 -0.0198 1.0000 1.0000
0.500 0.0973 0.01135 0.00294 -0.0190 1.0000 1.0000
0.750 0.1194 0.01149 0.00305 -0.0181 1.0000 1.0000
1.000 0.1415 0.01164 0.00321 -0.0172 1.0000 1.0000
1.250 0.1634 0.01181 0.00341 -0.0164 1.0000 1.0000
1.500 0.1853 0.01200 0.00364 -0.0155 1.0000 1.0000
1.750 0.2071 0.01220 0.00393 -0.0147 1.0000 1.0000
2.000 0.2487 0.01248 0.00442 -0.0181 0.9918 1.0000
2.250 0.3153 0.01258 0.00485 -0.0262 0.9693 1.0000
2.500 0.3919 0.01189 0.00471 -0.0350 0.9201 1.0000
2.750 0.4704 0.01421 0.00423 -0.0428 0.1335 1.0000
3.000 0.4896 0.01578 0.00530 -0.0410 0.0393 1.0000
3.250 0.5112 0.01684 0.00643 -0.0394 0.0269 1.0000
3.500 0.5323 0.01809 0.00788 -0.0376 0.0227 1.0000
3.750 0.5530 0.02010 0.00999 -0.0360 0.0188 1.0000
4.000 0.5782 0.02205 0.01219 -0.0348 0.0178 1.0000
4.250 0.6042 0.02441 0.01491 -0.0336 0.0171 1.0000
4.500 0.6289 0.02672 0.01764 -0.0322 0.0158 1.0000
4.750 0.6514 0.02916 0.02048 -0.0307 0.0146 1.0000
5.000 0.6727 0.03256 0.02450 -0.0286 0.0150 1.0000
5.250 0.6915 0.03626 0.02873 -0.0263 0.0157 1.0000
5.500 0.7078 0.04012 0.03305 -0.0242 0.0166 1.0000
5.750 0.7213 0.04417 0.03748 -0.0222 0.0176 1.0000
6.250 0.7513 0.05331 0.04749 -0.0182 0.0250 1.0000
6.500 0.7660 0.06013 0.05472 -0.0166 0.0515 1.0000
6.750 0.7740 0.06437 0.05918 -0.0160 0.0512 1.0000
7.000 0.7803 0.06851 0.06349 -0.0158 0.0500 1.0000
7.250 0.7848 0.07277 0.06786 -0.0155 0.0487 1.0000
7.500 0.7892 0.07691 0.07206 -0.0151 0.0475 1.0000
7.750 0.7969 0.08394 0.07887 -0.0143 0.0459 1.0000
8.000 0.7922 0.08725 0.08241 -0.0149 0.0457 1.0000
8.250 0.7839 0.09078 0.08609 -0.0172 0.0452 1.0000
8.500 0.7754 0.09472 0.09005 -0.0192 0.0449 1.0000
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