GOE 374 AIRFOIL (goe374-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 374 AIRFOIL (goe374-il) Reynolds number: 100,000 Max Cl/Cd: 51.76 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe374-il-100000.txt Download as CSV file: xf-goe374-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 374 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3441 0.09554 0.09079 -0.0162 1.0000 0.0360
-7.500 -0.3420 0.09301 0.08832 -0.0167 1.0000 0.0367
-7.250 -0.3413 0.09060 0.08599 -0.0172 1.0000 0.0375
-7.000 -0.3378 0.08799 0.08344 -0.0186 1.0000 0.0382
-6.750 -0.3325 0.08545 0.08096 -0.0208 1.0000 0.0391
-6.500 -0.3247 0.08314 0.07868 -0.0240 1.0000 0.0400
-6.250 -0.3120 0.08138 0.07691 -0.0292 1.0000 0.0407
-6.000 -0.2969 0.07970 0.07514 -0.0335 1.0000 0.0410
-5.750 -0.2909 0.07518 0.07069 -0.0333 1.0000 0.0416
-5.500 -0.2916 0.07056 0.06620 -0.0295 1.0000 0.0428
-5.250 -0.2871 0.06749 0.06318 -0.0278 1.0000 0.0448
-5.000 -0.2782 0.06480 0.06050 -0.0279 1.0000 0.0469
-4.750 -0.2653 0.06224 0.05790 -0.0291 1.0000 0.0503
-4.250 -0.2335 0.05666 0.05209 -0.0321 1.0000 0.0551
-4.000 -0.2271 0.05374 0.04922 -0.0306 1.0000 0.0575
-3.750 -0.2133 0.05129 0.04670 -0.0306 1.0000 0.0619
-3.500 -0.1835 0.05022 0.04518 -0.0331 1.0000 0.0674
-3.250 -0.1791 0.04621 0.04138 -0.0315 1.0000 0.0703
-3.000 -0.1643 0.04415 0.03925 -0.0312 1.0000 0.0762
-2.750 -0.1445 0.04201 0.03688 -0.0318 1.0000 0.0829
-2.500 -0.1038 0.03898 0.03365 -0.0363 0.9936 0.0967
-2.250 -0.0582 0.03582 0.03029 -0.0419 0.9841 0.1251
-2.000 -0.0178 0.03300 0.02742 -0.0463 0.9732 0.1710
-1.500 0.0485 0.02728 0.02196 -0.0509 0.9505 0.3542
-1.250 0.0844 0.02436 0.01924 -0.0518 0.9409 0.4471
-1.000 0.1433 0.02238 0.01681 -0.0575 0.9284 0.4345
-0.750 0.2371 0.02231 0.01472 -0.0646 0.9172 0.1628
-0.500 0.2866 0.02055 0.01253 -0.0665 0.9043 0.1282
-0.250 0.3323 0.01940 0.01097 -0.0679 0.8892 0.1097
0.000 0.3732 0.01816 0.00963 -0.0689 0.8721 0.1026
0.250 0.4075 0.01722 0.00860 -0.0688 0.8507 0.1006
0.500 0.4392 0.01643 0.00775 -0.0683 0.8279 0.1038
0.750 0.4663 0.01584 0.00709 -0.0669 0.8014 0.1031
1.000 0.4930 0.01546 0.00656 -0.0657 0.7739 0.1037
1.250 0.5194 0.01524 0.00613 -0.0644 0.7458 0.1059
1.500 0.5457 0.01509 0.00578 -0.0633 0.7184 0.1109
1.750 0.5718 0.01503 0.00555 -0.0623 0.6924 0.1254
2.000 0.6180 0.01328 0.00546 -0.0657 0.6659 1.0000
2.250 0.6430 0.01363 0.00547 -0.0646 0.6438 1.0000
2.500 0.6677 0.01402 0.00560 -0.0637 0.6231 1.0000
2.750 0.6924 0.01442 0.00582 -0.0628 0.6038 1.0000
3.000 0.7173 0.01482 0.00604 -0.0620 0.5864 1.0000
3.250 0.7422 0.01521 0.00626 -0.0613 0.5698 1.0000
3.500 0.7663 0.01556 0.00654 -0.0604 0.5518 1.0000
3.750 0.7903 0.01587 0.00680 -0.0595 0.5336 1.0000
4.000 0.8142 0.01613 0.00695 -0.0585 0.5154 1.0000
4.250 0.8383 0.01643 0.00712 -0.0576 0.4984 1.0000
4.500 0.8623 0.01678 0.00744 -0.0568 0.4818 1.0000
4.750 0.8865 0.01717 0.00781 -0.0560 0.4664 1.0000
5.000 0.9106 0.01760 0.00826 -0.0553 0.4514 1.0000
5.250 0.9347 0.01806 0.00872 -0.0546 0.4365 1.0000
5.500 0.9586 0.01854 0.00921 -0.0538 0.4210 1.0000
5.750 0.9821 0.01902 0.00970 -0.0530 0.4048 1.0000
6.000 1.0054 0.01950 0.01023 -0.0521 0.3881 1.0000
6.250 1.0286 0.01997 0.01070 -0.0512 0.3712 1.0000
6.500 1.0510 0.02047 0.01128 -0.0502 0.3532 1.0000
6.750 1.0729 0.02098 0.01188 -0.0492 0.3340 1.0000
7.000 1.0949 0.02144 0.01232 -0.0481 0.3149 1.0000
7.250 1.1146 0.02175 0.01276 -0.0467 0.2924 1.0000
7.500 1.1342 0.02202 0.01295 -0.0454 0.2716 1.0000
7.750 1.1529 0.02237 0.01344 -0.0440 0.2498 1.0000
8.000 1.1720 0.02281 0.01390 -0.0428 0.2315 1.0000
8.250 1.1919 0.02341 0.01461 -0.0416 0.2171 1.0000
8.500 1.2103 0.02389 0.01527 -0.0404 0.2000 1.0000
8.750 1.2283 0.02429 0.01586 -0.0391 0.1786 1.0000
9.000 1.2447 0.02490 0.01652 -0.0377 0.1494 1.0000
9.250 1.2583 0.02613 0.01768 -0.0360 0.1095 1.0000
9.500 1.2650 0.02825 0.01945 -0.0336 0.0601 1.0000
9.750 1.2683 0.03048 0.02155 -0.0308 0.0487 1.0000
10.000 1.2742 0.03226 0.02352 -0.0282 0.0433 1.0000
10.250 1.2741 0.03430 0.02570 -0.0252 0.0400 1.0000
10.500 1.2679 0.03654 0.02802 -0.0216 0.0384 1.0000
10.750 1.2671 0.03854 0.03022 -0.0190 0.0376 1.0000
11.000 1.2660 0.04077 0.03262 -0.0167 0.0368 1.0000
11.250 1.2657 0.04320 0.03521 -0.0148 0.0360 1.0000
11.500 1.2675 0.04573 0.03789 -0.0132 0.0354 1.0000
11.750 1.2714 0.04837 0.04069 -0.0118 0.0348 1.0000
12.000 1.2770 0.05123 0.04372 -0.0104 0.0343 1.0000
12.250 1.2822 0.05424 0.04694 -0.0093 0.0341 1.0000
12.500 1.2817 0.05764 0.05057 -0.0087 0.0333 1.0000
12.750 1.2784 0.06127 0.05441 -0.0085 0.0327 1.0000
13.000 1.2730 0.06533 0.05872 -0.0087 0.0326 1.0000
13.250 1.2645 0.06968 0.06329 -0.0095 0.0318 1.0000
13.500 1.2531 0.07472 0.06860 -0.0110 0.0320 1.0000
13.750 1.2404 0.08000 0.07408 -0.0131 0.0316 1.0000
14.000 1.2217 0.08665 0.08102 -0.0164 0.0322 1.0000
14.250 1.1979 0.09467 0.08931 -0.0209 0.0334 1.0000
14.500 1.1748 0.10287 0.09772 -0.0261 0.0337 1.0000
14.750 1.1454 0.11321 0.10828 -0.0329 0.0351 1.0000
15.000 1.1179 0.12389 0.11911 -0.0398 0.0361 1.0000
15.250 1.0923 0.13502 0.13032 -0.0469 0.0372 1.0000
15.500 1.0694 0.14627 0.14160 -0.0536 0.0380 1.0000
15.750 1.0529 0.15656 0.15191 -0.0594 0.0387 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 374 AIRFOIL (goe374-il)