GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 500,000 Max Cl/Cd: 88.55 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe622-il-500000.txt Download as CSV file: xf-goe622-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 622 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.3964 0.11231 0.11011 -0.0120 1.0000 0.0192
-10.500 -0.3963 0.10837 0.10617 -0.0132 1.0000 0.0200
-9.000 -0.5065 0.09458 0.09236 -0.0126 1.0000 0.0214
-7.250 -0.5318 0.04026 0.03759 -0.0463 1.0000 0.0152
-7.000 -0.5386 0.03097 0.02767 -0.0461 1.0000 0.0144
-6.750 -0.5332 0.02526 0.02132 -0.0443 1.0000 0.0142
-6.500 -0.5195 0.02261 0.01826 -0.0425 1.0000 0.0148
-6.250 -0.5031 0.02122 0.01659 -0.0407 1.0000 0.0153
-6.000 -0.4788 0.01936 0.01437 -0.0405 0.9991 0.0157
-5.750 -0.4490 0.01561 0.01011 -0.0417 0.9973 0.0167
-5.500 -0.4148 0.01453 0.00890 -0.0434 0.9953 0.0178
-5.250 -0.3806 0.01370 0.00795 -0.0448 0.9931 0.0192
-5.000 -0.3466 0.01306 0.00720 -0.0462 0.9901 0.0211
-4.750 -0.3129 0.01192 0.00593 -0.0476 0.9874 0.0237
-4.500 -0.2775 0.01131 0.00529 -0.0493 0.9850 0.0269
-4.250 -0.2416 0.01093 0.00482 -0.0510 0.9827 0.0303
-4.000 -0.2096 0.01019 0.00408 -0.0520 0.9780 0.0383
-3.750 -0.1744 0.00966 0.00353 -0.0536 0.9746 0.0473
-3.500 -0.1388 0.00926 0.00313 -0.0553 0.9717 0.0594
-3.250 -0.1066 0.00894 0.00288 -0.0562 0.9662 0.0794
-3.000 -0.0738 0.00871 0.00273 -0.0572 0.9603 0.1045
-2.750 -0.0411 0.00850 0.00253 -0.0582 0.9545 0.1217
-2.500 -0.0109 0.00829 0.00233 -0.0587 0.9466 0.1347
-2.250 0.0190 0.00808 0.00215 -0.0590 0.9385 0.1498
-2.000 0.0481 0.00784 0.00199 -0.0592 0.9299 0.1733
-1.750 0.0751 0.00754 0.00185 -0.0590 0.9196 0.2244
-1.500 0.1012 0.00703 0.00175 -0.0587 0.9096 0.3534
-1.250 0.1247 0.00624 0.00170 -0.0581 0.8992 0.5790
-1.000 0.1441 0.00553 0.00173 -0.0559 0.8872 0.7994
-0.750 0.1655 0.00530 0.00175 -0.0535 0.8734 0.9209
-0.500 0.2050 0.00529 0.00168 -0.0556 0.8571 0.9757
-0.250 0.2477 0.00530 0.00158 -0.0586 0.8355 0.9979
0.000 0.2754 0.00534 0.00149 -0.0584 0.8147 1.0000
0.250 0.3009 0.00538 0.00144 -0.0578 0.7941 1.0000
0.500 0.3265 0.00545 0.00139 -0.0571 0.7735 1.0000
0.750 0.3522 0.00552 0.00136 -0.0565 0.7500 1.0000
1.000 0.3777 0.00563 0.00135 -0.0559 0.7233 1.0000
1.250 0.4032 0.00576 0.00134 -0.0553 0.6954 1.0000
1.500 0.4288 0.00592 0.00136 -0.0548 0.6694 1.0000
1.750 0.4544 0.00609 0.00141 -0.0542 0.6445 1.0000
2.000 0.4805 0.00625 0.00148 -0.0538 0.6225 1.0000
2.250 0.5065 0.00643 0.00155 -0.0534 0.6021 1.0000
2.500 0.5329 0.00658 0.00164 -0.0531 0.5827 1.0000
2.750 0.5589 0.00677 0.00173 -0.0527 0.5608 1.0000
3.000 0.5852 0.00694 0.00185 -0.0523 0.5385 1.0000
3.250 0.6116 0.00712 0.00196 -0.0520 0.5183 1.0000
3.500 0.6380 0.00729 0.00208 -0.0517 0.4950 1.0000
3.750 0.6641 0.00750 0.00222 -0.0514 0.4663 1.0000
4.000 0.6893 0.00780 0.00236 -0.0509 0.4201 1.0000
4.250 0.7133 0.00828 0.00256 -0.0503 0.3557 1.0000
4.500 0.7370 0.00885 0.00285 -0.0498 0.2944 1.0000
4.750 0.7608 0.00943 0.00318 -0.0493 0.2399 1.0000
5.000 0.7855 0.00988 0.00347 -0.0489 0.2025 1.0000
5.250 0.8099 0.01039 0.00379 -0.0484 0.1582 1.0000
5.500 0.8340 0.01093 0.00416 -0.0480 0.1268 1.0000
5.750 0.8589 0.01137 0.00452 -0.0476 0.1062 1.0000
6.000 0.8814 0.01214 0.00500 -0.0469 0.0541 1.0000
6.250 0.9033 0.01305 0.00570 -0.0460 0.0233 1.0000
6.500 0.9276 0.01361 0.00634 -0.0453 0.0192 1.0000
6.750 0.9513 0.01422 0.00704 -0.0446 0.0175 1.0000
7.000 0.9740 0.01498 0.00788 -0.0438 0.0160 1.0000
7.250 0.9943 0.01608 0.00912 -0.0425 0.0149 1.0000
7.500 1.0129 0.01739 0.01055 -0.0410 0.0142 1.0000
7.750 1.0348 0.01817 0.01142 -0.0401 0.0137 1.0000
8.000 1.0562 0.01902 0.01236 -0.0392 0.0129 1.0000
8.250 1.0761 0.02009 0.01352 -0.0380 0.0124 1.0000
8.500 1.0953 0.02130 0.01484 -0.0367 0.0119 1.0000
8.750 1.1140 0.02261 0.01627 -0.0354 0.0116 1.0000
9.000 1.1323 0.02405 0.01783 -0.0341 0.0112 1.0000
9.250 1.1498 0.02567 0.01957 -0.0328 0.0109 1.0000
9.500 1.1663 0.02755 0.02161 -0.0313 0.0107 1.0000
9.750 1.1816 0.02973 0.02398 -0.0297 0.0106 1.0000
10.000 1.1946 0.03233 0.02689 -0.0278 0.0108 1.0000
10.250 1.1903 0.03859 0.03392 -0.0237 0.0127 1.0000
10.750 1.1487 0.05334 0.04983 -0.0147 0.0197 1.0000
11.000 1.1391 0.05528 0.05197 -0.0117 0.0190 1.0000
11.250 1.1237 0.05849 0.05538 -0.0099 0.0186 1.0000
11.500 1.1059 0.06250 0.05957 -0.0095 0.0184 1.0000
11.750 1.0849 0.06764 0.06489 -0.0109 0.0185 1.0000
12.000 1.0623 0.07394 0.07136 -0.0141 0.0187 1.0000
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Polar data table (+)
Polar graphs
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