GOE 622 AIRFOIL (goe622-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 622 AIRFOIL (goe622-il) Reynolds number: 100,000 Max Cl/Cd: 53.25 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe622-il-100000.txt Download as CSV file: xf-goe622-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4956 0.10711 0.10216 -0.0099 1.0000 0.0897
-8.750 -0.5095 0.10477 0.09994 -0.0170 1.0000 0.0912
-8.500 -0.5215 0.10140 0.09669 -0.0257 1.0000 0.0916
-8.250 -0.4863 0.09550 0.09068 -0.0136 1.0000 0.0958
-8.000 -0.4825 0.09228 0.08750 -0.0147 1.0000 0.1005
-7.750 -0.4909 0.08887 0.08420 -0.0208 1.0000 0.1041
-7.500 -0.5010 0.08437 0.07959 -0.0353 1.0000 0.1056
-7.250 -0.4801 0.08062 0.07602 -0.0229 1.0000 0.1103
-7.000 -0.4741 0.07685 0.07226 -0.0256 1.0000 0.1154
-6.500 -0.4638 0.06797 0.06335 -0.0322 1.0000 0.1237
-6.250 -0.4544 0.06451 0.05987 -0.0328 1.0000 0.1291
-6.000 -0.4481 0.05979 0.05500 -0.0367 1.0000 0.1364
-5.750 -0.4363 0.05668 0.05191 -0.0354 1.0000 0.1406
-5.500 -0.4176 0.04240 0.03624 -0.0430 1.0000 0.0776
-5.250 -0.4027 0.03824 0.03203 -0.0422 1.0000 0.0742
-5.000 -0.3855 0.03424 0.02755 -0.0415 1.0000 0.0739
-4.750 -0.3661 0.03059 0.02336 -0.0406 1.0000 0.0739
-4.500 -0.3444 0.02728 0.01945 -0.0396 1.0000 0.0739
-4.250 -0.3209 0.02483 0.01635 -0.0385 1.0000 0.0771
-4.000 -0.2989 0.02329 0.01474 -0.0378 1.0000 0.0847
-3.750 -0.2737 0.02177 0.01266 -0.0368 1.0000 0.0916
-3.500 -0.2502 0.02059 0.01144 -0.0362 1.0000 0.1044
-3.250 -0.2255 0.01936 0.01014 -0.0356 1.0000 0.1181
-3.000 -0.2005 0.01825 0.00900 -0.0352 1.0000 0.1364
-2.750 -0.1755 0.01735 0.00814 -0.0348 1.0000 0.1613
-2.500 -0.1500 0.01645 0.00741 -0.0346 1.0000 0.1932
-2.250 -0.1247 0.01578 0.00694 -0.0345 1.0000 0.2313
-2.000 -0.0992 0.01515 0.00657 -0.0344 1.0000 0.2798
-1.750 -0.0718 0.01374 0.00629 -0.0349 1.0000 0.4788
-1.500 -0.0489 0.01245 0.00626 -0.0320 1.0000 1.0000
-1.250 -0.0267 0.01266 0.00619 -0.0316 1.0000 1.0000
-1.000 -0.0047 0.01291 0.00621 -0.0312 1.0000 1.0000
-0.750 0.0170 0.01321 0.00634 -0.0309 1.0000 1.0000
-0.500 0.0383 0.01358 0.00657 -0.0307 1.0000 1.0000
-0.250 0.0883 0.01394 0.00676 -0.0359 0.9898 1.0000
0.000 0.1388 0.01426 0.00693 -0.0410 0.9784 1.0000
0.250 0.1886 0.01451 0.00709 -0.0459 0.9667 1.0000
0.500 0.2385 0.01468 0.00719 -0.0507 0.9547 1.0000
0.750 0.2893 0.01475 0.00723 -0.0554 0.9431 1.0000
1.000 0.3382 0.01473 0.00720 -0.0596 0.9298 1.0000
1.250 0.3864 0.01455 0.00704 -0.0631 0.9154 1.0000
1.500 0.4272 0.01436 0.00686 -0.0649 0.8980 1.0000
1.750 0.4622 0.01424 0.00676 -0.0655 0.8797 1.0000
2.000 0.4944 0.01416 0.00672 -0.0656 0.8625 1.0000
2.250 0.5245 0.01407 0.00664 -0.0651 0.8449 1.0000
2.500 0.5502 0.01406 0.00664 -0.0638 0.8244 1.0000
2.750 0.5760 0.01399 0.00658 -0.0624 0.8037 1.0000
3.000 0.6007 0.01395 0.00657 -0.0609 0.7816 1.0000
3.250 0.6258 0.01390 0.00651 -0.0594 0.7596 1.0000
3.500 0.6500 0.01393 0.00655 -0.0579 0.7352 1.0000
3.750 0.6747 0.01397 0.00662 -0.0565 0.7111 1.0000
4.000 0.7000 0.01406 0.00670 -0.0553 0.6877 1.0000
4.250 0.7244 0.01423 0.00688 -0.0540 0.6612 1.0000
4.500 0.7486 0.01444 0.00706 -0.0527 0.6323 1.0000
4.750 0.7721 0.01470 0.00733 -0.0513 0.5991 1.0000
5.000 0.7949 0.01499 0.00758 -0.0497 0.5612 1.0000
5.250 0.8168 0.01534 0.00785 -0.0481 0.5171 1.0000
5.500 0.8375 0.01573 0.00816 -0.0463 0.4617 1.0000
5.750 0.8568 0.01628 0.00849 -0.0445 0.3945 1.0000
6.000 0.8745 0.01717 0.00901 -0.0427 0.3261 1.0000
6.250 0.8929 0.01818 0.00972 -0.0413 0.2735 1.0000
6.500 0.9126 0.01913 0.01051 -0.0401 0.2343 1.0000
6.750 0.9326 0.01991 0.01123 -0.0391 0.1880 1.0000
7.000 0.9476 0.02152 0.01247 -0.0375 0.1198 1.0000
7.250 0.9602 0.02382 0.01437 -0.0350 0.0734 1.0000
7.500 0.9759 0.02563 0.01606 -0.0333 0.0608 1.0000
7.750 0.9955 0.02736 0.01795 -0.0316 0.0551 1.0000
8.000 1.0155 0.02939 0.01997 -0.0303 0.0511 1.0000
8.250 1.0382 0.03254 0.02315 -0.0294 0.0485 1.0000
8.500 1.0597 0.03456 0.02556 -0.0281 0.0460 1.0000
8.750 1.0792 0.03696 0.02828 -0.0268 0.0435 1.0000
9.000 1.0965 0.04011 0.03184 -0.0253 0.0428 1.0000
9.250 1.1090 0.04375 0.03602 -0.0234 0.0428 1.0000
9.500 1.1154 0.04789 0.04071 -0.0211 0.0433 1.0000
9.750 1.1159 0.05240 0.04575 -0.0187 0.0441 1.0000
10.000 1.1111 0.05704 0.05083 -0.0163 0.0451 1.0000
10.250 1.1021 0.06164 0.05578 -0.0141 0.0460 1.0000
10.500 1.0893 0.06610 0.06050 -0.0121 0.0468 1.0000
10.750 1.0729 0.07035 0.06495 -0.0102 0.0474 1.0000
11.000 1.0564 0.07510 0.06984 -0.0094 0.0481 1.0000
11.250 1.0515 0.07965 0.07452 -0.0092 0.0496 1.0000
11.500 1.0194 0.08498 0.08013 -0.0118 0.0503 1.0000
11.750 0.8759 0.08249 0.07780 -0.0041 0.0478 1.0000
12.000 0.8503 0.08963 0.08501 -0.0071 0.0483 1.0000
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Polar data table (+)
Polar graphs
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