GOE 165 (MVA MK.11) AIRFOIL (goe165-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 165 (MVA MK.11) AIRFOIL (goe165-il) Reynolds number: 200,000 Max Cl/Cd: 69.3 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe165-il-200000.txt Download as CSV file: xf-goe165-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 165 (MVA MK.11) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3265 0.10288 0.09970 -0.0361 1.0000 0.0487
-8.500 -0.3353 0.09998 0.09687 -0.0368 1.0000 0.0489
-8.250 -0.3457 0.09718 0.09415 -0.0367 1.0000 0.0490
-8.000 -0.3329 0.09289 0.08987 -0.0331 1.0000 0.0507
-7.750 -0.3314 0.09098 0.08800 -0.0311 1.0000 0.0517
-7.500 -0.3356 0.08907 0.08616 -0.0292 1.0000 0.0526
-7.250 -0.3452 0.08731 0.08448 -0.0268 1.0000 0.0535
-7.000 -0.3576 0.08557 0.08282 -0.0245 1.0000 0.0543
-6.750 -0.3669 0.08330 0.08063 -0.0236 1.0000 0.0553
-6.500 -0.3260 0.06192 0.05916 -0.0655 0.9880 0.0621
-6.250 -0.2988 0.06401 0.06132 -0.0579 0.9850 0.0639
-6.000 -0.2331 0.04525 0.04205 -0.0974 0.9750 0.0755
-5.500 -0.1120 0.02854 0.02318 -0.1210 0.9647 0.0366
-5.250 -0.0651 0.02407 0.01826 -0.1260 0.9614 0.0330
-5.000 -0.0232 0.02149 0.01528 -0.1287 0.9538 0.0316
-4.750 0.0210 0.02002 0.01365 -0.1319 0.9481 0.0341
-4.500 0.0590 0.01855 0.01206 -0.1335 0.9383 0.0350
-4.250 0.1000 0.01716 0.01061 -0.1357 0.9311 0.0358
-4.000 0.1394 0.01609 0.00949 -0.1375 0.9220 0.0374
-3.750 0.1797 0.01470 0.00811 -0.1402 0.9114 0.0417
-3.500 0.2256 0.01356 0.00679 -0.1438 0.9021 0.0549
-3.250 0.2754 0.01178 0.00584 -0.1492 0.8923 0.3058
-3.000 0.3181 0.01204 0.00599 -0.1514 0.8764 0.3466
-2.750 0.3584 0.01218 0.00597 -0.1533 0.8554 0.3607
-2.500 0.3963 0.01277 0.00642 -0.1546 0.8342 0.3811
-2.250 0.4290 0.01306 0.00663 -0.1550 0.8133 0.3887
-2.000 0.4610 0.01316 0.00646 -0.1556 0.7935 0.3977
-1.750 0.4902 0.01335 0.00657 -0.1554 0.7754 0.4022
-1.500 0.5188 0.01350 0.00654 -0.1552 0.7584 0.4087
-1.250 0.5467 0.01358 0.00646 -0.1550 0.7423 0.4142
-1.000 0.5737 0.01373 0.00651 -0.1545 0.7274 0.4185
-0.750 0.6012 0.01387 0.00648 -0.1543 0.7127 0.4256
-0.500 0.6271 0.01405 0.00663 -0.1536 0.6980 0.4314
-0.250 0.6533 0.01426 0.00673 -0.1530 0.6823 0.4398
0.000 0.6786 0.01439 0.00682 -0.1521 0.6653 0.4455
0.250 0.7039 0.01451 0.00686 -0.1514 0.6467 0.4513
0.500 0.7293 0.01460 0.00681 -0.1507 0.6260 0.4574
0.750 0.7530 0.01467 0.00684 -0.1496 0.6004 0.4618
1.000 0.7764 0.01477 0.00679 -0.1485 0.5667 0.4674
1.250 0.7989 0.01487 0.00668 -0.1473 0.5204 0.4720
1.500 0.8200 0.01514 0.00669 -0.1458 0.4757 0.4760
1.750 0.8425 0.01551 0.00679 -0.1448 0.4469 0.4812
2.000 0.8667 0.01582 0.00691 -0.1441 0.4279 0.4860
2.250 0.8911 0.01609 0.00709 -0.1435 0.4145 0.4901
2.500 0.9162 0.01640 0.00728 -0.1430 0.4039 0.4942
2.750 0.9428 0.01660 0.00743 -0.1428 0.3947 0.4978
3.000 0.9685 0.01683 0.00759 -0.1425 0.3877 0.5008
3.250 0.9947 0.01699 0.00778 -0.1422 0.3808 0.5041
3.500 1.0207 0.01727 0.00801 -0.1419 0.3749 0.5072
3.750 1.0471 0.01744 0.00821 -0.1417 0.3684 0.5104
4.000 1.0726 0.01769 0.00840 -0.1413 0.3609 0.5136
4.250 1.0980 0.01783 0.00861 -0.1409 0.3538 0.5168
4.500 1.1232 0.01802 0.00883 -0.1405 0.3472 0.5199
4.750 1.1486 0.01824 0.00908 -0.1401 0.3410 0.5231
5.000 1.1740 0.01842 0.00930 -0.1396 0.3349 0.5264
5.250 1.1990 0.01868 0.00956 -0.1392 0.3300 0.5296
5.500 1.2229 0.01875 0.00979 -0.1385 0.3216 0.5334
5.750 1.2457 0.01890 0.00996 -0.1376 0.3108 0.5375
6.000 1.2692 0.01907 0.01016 -0.1369 0.3017 0.5413
6.250 1.2931 0.01918 0.01041 -0.1362 0.2940 0.5448
6.500 1.3156 0.01934 0.01063 -0.1353 0.2831 0.5490
6.750 1.3381 0.01946 0.01081 -0.1344 0.2669 0.5536
7.000 1.3608 0.01965 0.01105 -0.1335 0.2518 0.5584
7.250 1.3818 0.01994 0.01132 -0.1324 0.2280 0.5640
7.500 1.3961 0.02084 0.01186 -0.1304 0.1655 0.5695
7.750 1.4045 0.02242 0.01310 -0.1277 0.1155 0.5742
8.000 1.4196 0.02334 0.01400 -0.1257 0.0941 0.5804
8.250 1.4267 0.02483 0.01531 -0.1226 0.0555 0.5863
8.500 1.4404 0.02559 0.01625 -0.1203 0.0483 0.5939
8.750 1.4518 0.02645 0.01727 -0.1177 0.0411 0.6029
9.000 1.4656 0.02720 0.01817 -0.1154 0.0281 0.6161
9.250 1.4745 0.02827 0.01929 -0.1128 0.0237 0.6392
9.500 1.4785 0.02889 0.02041 -0.1090 0.0221 1.0000
9.750 1.4855 0.03020 0.02183 -0.1062 0.0210 1.0000
10.000 1.4902 0.03169 0.02347 -0.1033 0.0203 1.0000
10.250 1.4919 0.03347 0.02542 -0.1003 0.0198 1.0000
10.500 1.4916 0.03548 0.02762 -0.0976 0.0195 1.0000
10.750 1.4891 0.03782 0.03014 -0.0951 0.0194 1.0000
11.000 1.4845 0.04054 0.03304 -0.0929 0.0192 1.0000
11.250 1.4780 0.04370 0.03637 -0.0914 0.0190 1.0000
11.500 1.4704 0.04728 0.04012 -0.0904 0.0188 1.0000
11.750 1.4618 0.05126 0.04426 -0.0899 0.0186 1.0000
12.000 1.4524 0.05555 0.04870 -0.0898 0.0183 1.0000
12.250 1.4425 0.05999 0.05327 -0.0899 0.0181 1.0000
12.500 1.4323 0.06454 0.05796 -0.0901 0.0178 1.0000
12.750 1.4225 0.06910 0.06264 -0.0903 0.0176 1.0000
13.000 1.4141 0.07348 0.06712 -0.0904 0.0175 1.0000
13.250 1.4072 0.07766 0.07140 -0.0905 0.0175 1.0000
13.500 1.4025 0.08148 0.07532 -0.0903 0.0174 1.0000
13.750 1.4007 0.08477 0.07871 -0.0898 0.0174 1.0000
14.000 1.4021 0.08749 0.08150 -0.0887 0.0174 1.0000
14.250 1.4064 0.08975 0.08384 -0.0872 0.0175 1.0000
14.500 1.4128 0.09171 0.08590 -0.0854 0.0176 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 165 (MVA MK.11) AIRFOIL (goe165-il)