GOE 368 AIRFOIL (goe368-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 368 AIRFOIL (goe368-il) Reynolds number: 50,000 Max Cl/Cd: 29.27 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe368-il-50000.txt Download as CSV file: xf-goe368-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 368 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.4405 0.13066 0.12378 0.0076 1.0000 0.1417
-9.750 -0.4426 0.12937 0.12257 0.0054 1.0000 0.1448
-9.500 -0.4529 0.12985 0.12317 0.0023 1.0000 0.1458
-9.250 -0.4269 0.12118 0.11447 0.0041 1.0000 0.1492
-9.000 -0.4185 0.11758 0.11089 0.0036 1.0000 0.1532
-8.750 -0.4169 0.11529 0.10868 0.0021 1.0000 0.1573
-8.500 -0.4261 0.11509 0.10860 -0.0007 1.0000 0.1593
-8.250 -0.4132 0.10954 0.10310 -0.0006 1.0000 0.1615
-8.000 -0.4000 0.10511 0.09868 -0.0002 1.0000 0.1657
-7.750 -0.3980 0.10263 0.09627 -0.0013 1.0000 0.1705
-7.500 -0.4087 0.10203 0.09583 -0.0038 1.0000 0.1731
-7.250 -0.3935 0.09660 0.09043 -0.0029 1.0000 0.1771
-7.000 -0.3873 0.09349 0.08738 -0.0030 1.0000 0.1828
-6.750 -0.3936 0.09243 0.08645 -0.0081 1.0000 0.1873
-6.500 -0.3812 0.08751 0.08154 -0.0067 1.0000 0.1907
-6.250 -0.3721 0.08410 0.07818 -0.0071 1.0000 0.1964
-6.000 -0.3706 0.08240 0.07655 -0.0141 1.0000 0.2020
-5.750 -0.3576 0.07752 0.07174 -0.0114 1.0000 0.2052
-5.500 -0.3472 0.07430 0.06855 -0.0129 1.0000 0.2119
-5.250 -0.3369 0.07095 0.06522 -0.0164 1.0000 0.2173
-5.000 -0.3243 0.06727 0.06159 -0.0160 1.0000 0.2225
-4.750 -0.3097 0.06408 0.05836 -0.0203 1.0000 0.2306
-4.500 -0.2959 0.06035 0.05468 -0.0198 1.0000 0.2352
-4.250 -0.2470 0.04998 0.04359 -0.0357 1.0000 0.1285
-4.000 -0.2242 0.04557 0.03898 -0.0377 1.0000 0.1250
-3.750 -0.1988 0.04100 0.03407 -0.0400 1.0000 0.1229
-3.500 -0.1728 0.03661 0.02925 -0.0416 1.0000 0.1222
-3.250 -0.1480 0.03321 0.02548 -0.0420 1.0000 0.1255
-3.000 -0.1223 0.03001 0.02173 -0.0423 1.0000 0.1347
-2.750 -0.1003 0.02868 0.02018 -0.0414 1.0000 0.1579
-2.500 -0.0786 0.02677 0.01807 -0.0402 1.0000 0.2031
-2.250 -0.0651 0.02578 0.01721 -0.0382 1.0000 0.2759
-2.000 -0.0553 0.02539 0.01701 -0.0358 1.0000 0.3424
-1.750 -0.0355 0.02465 0.01598 -0.0353 1.0000 0.3543
-1.500 -0.0153 0.02408 0.01514 -0.0350 1.0000 0.3573
-1.250 0.0032 0.02373 0.01462 -0.0346 1.0000 0.3626
-1.000 0.0222 0.02361 0.01428 -0.0345 1.0000 0.3663
-0.750 0.0405 0.02364 0.01410 -0.0344 1.0000 0.3684
-0.500 0.0884 0.02363 0.01386 -0.0397 0.9858 0.3736
-0.250 0.1652 0.02335 0.01333 -0.0494 0.9570 0.3876
0.000 0.2404 0.02271 0.01264 -0.0581 0.9275 0.4105
0.250 0.3115 0.02168 0.01178 -0.0651 0.8974 0.4414
0.500 0.3646 0.02023 0.01105 -0.0684 0.8657 0.5221
0.750 0.4287 0.01889 0.01025 -0.0720 0.8304 1.0000
1.000 0.4605 0.01912 0.01006 -0.0707 0.7933 1.0000
1.250 0.4872 0.01941 0.01000 -0.0686 0.7587 1.0000
1.500 0.5104 0.01979 0.01012 -0.0664 0.7237 1.0000
1.750 0.5340 0.02012 0.01020 -0.0642 0.6918 1.0000
2.000 0.5577 0.02049 0.01032 -0.0622 0.6619 1.0000
2.250 0.5813 0.02096 0.01057 -0.0604 0.6332 1.0000
2.500 0.6047 0.02152 0.01094 -0.0588 0.6051 1.0000
2.750 0.6282 0.02214 0.01144 -0.0573 0.5781 1.0000
3.000 0.6518 0.02278 0.01195 -0.0560 0.5528 1.0000
3.250 0.6761 0.02334 0.01234 -0.0547 0.5306 1.0000
3.500 0.6998 0.02407 0.01304 -0.0537 0.5082 1.0000
3.750 0.7247 0.02476 0.01361 -0.0527 0.4904 1.0000
4.000 0.7489 0.02561 0.01448 -0.0519 0.4733 1.0000
4.250 0.7729 0.02657 0.01545 -0.0512 0.4573 1.0000
4.500 0.7971 0.02758 0.01648 -0.0505 0.4436 1.0000
4.750 0.8222 0.02860 0.01746 -0.0498 0.4329 1.0000
5.000 0.8446 0.03010 0.01926 -0.0496 0.4223 1.0000
5.250 0.8669 0.03166 0.02101 -0.0492 0.4132 1.0000
5.500 0.8916 0.03284 0.02223 -0.0486 0.4054 1.0000
5.750 0.9103 0.03492 0.02467 -0.0485 0.3970 1.0000
6.000 0.9344 0.03631 0.02615 -0.0480 0.3912 1.0000
6.250 0.9498 0.03915 0.02947 -0.0482 0.3861 1.0000
6.500 0.9653 0.04198 0.03264 -0.0483 0.3817 1.0000
6.750 0.9868 0.04381 0.03464 -0.0479 0.3772 1.0000
7.000 1.0112 0.04366 0.03451 -0.0460 0.3613 1.0000
7.250 1.0301 0.04414 0.03516 -0.0445 0.3452 1.0000
7.500 1.0520 0.04418 0.03539 -0.0428 0.3299 1.0000
7.750 0.9132 0.07348 0.06518 -0.0603 0.3824 1.0000
8.000 0.8689 0.08386 0.07534 -0.0666 0.4064 1.0000
8.250 0.9578 0.07408 0.06613 -0.0553 0.3518 1.0000
8.500 0.9629 0.07610 0.06826 -0.0541 0.3388 1.0000
8.750 0.9579 0.08080 0.07300 -0.0553 0.3360 1.0000
9.000 0.9527 0.08580 0.07805 -0.0568 0.3348 1.0000
9.250 0.9467 0.09099 0.08330 -0.0585 0.3341 1.0000
9.500 0.9391 0.09627 0.08860 -0.0603 0.3335 1.0000
9.750 0.9294 0.10168 0.09404 -0.0624 0.3334 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 368 AIRFOIL (goe368-il)