GOE 137 (MVA H.15) AIRFOIL (goe137-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 137 (MVA H.15) AIRFOIL (goe137-il) Reynolds number: 200,000 Max Cl/Cd: 78.36 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe137-il-200000.txt Download as CSV file: xf-goe137-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 137 (MVA H.15) 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3442 0.09320 0.08988 -0.0236 1.0000 0.0387
-7.750 -0.3462 0.09082 0.08758 -0.0276 1.0000 0.0397
-7.500 -0.3420 0.08791 0.08472 -0.0339 1.0000 0.0402
-7.250 -0.3372 0.08475 0.08159 -0.0378 1.0000 0.0404
-7.000 -0.3334 0.08162 0.07847 -0.0400 1.0000 0.0406
-6.750 -0.3400 0.07630 0.07323 -0.0394 1.0000 0.0412
-6.500 -0.3425 0.07357 0.07056 -0.0358 1.0000 0.0419
-6.250 -0.3442 0.07160 0.06864 -0.0331 1.0000 0.0427
-6.000 -0.3375 0.06915 0.06621 -0.0329 0.9986 0.0438
-5.750 -0.3001 0.06453 0.06151 -0.0408 0.9932 0.0481
-5.500 -0.1754 0.03920 0.03619 -0.0630 0.9700 0.0572
-5.250 -0.1442 0.03452 0.03137 -0.0691 0.9606 0.0629
-5.000 -0.1556 0.04695 0.04296 -0.0732 0.9725 0.0664
-4.750 -0.1287 0.04278 0.03900 -0.0755 0.9676 0.0694
-4.500 -0.0554 0.02188 0.01821 -0.0812 0.9299 0.0839
-4.250 -0.0281 0.01859 0.01458 -0.0834 0.9160 0.0950
-4.000 -0.0070 0.01679 0.01276 -0.0829 0.9014 0.1003
-3.750 0.0160 0.01459 0.01033 -0.0832 0.8866 0.1118
-3.500 0.0398 0.01267 0.00812 -0.0833 0.8719 0.1246
-3.250 0.0630 0.01128 0.00659 -0.0828 0.8576 0.1319
-3.000 0.1090 0.01909 0.01232 -0.0881 0.8847 0.0657
-2.750 0.1353 0.01724 0.01016 -0.0873 0.8694 0.0648
-2.500 0.1618 0.01602 0.00866 -0.0864 0.8540 0.0646
-2.250 0.1883 0.01517 0.00757 -0.0856 0.8387 0.0652
-2.000 0.2147 0.01435 0.00657 -0.0849 0.8236 0.0671
-1.750 0.2411 0.01357 0.00570 -0.0842 0.8093 0.0684
-1.500 0.2676 0.01304 0.00511 -0.0836 0.7955 0.0703
-1.250 0.2945 0.01266 0.00466 -0.0831 0.7820 0.0732
-1.000 0.3215 0.01238 0.00431 -0.0826 0.7689 0.0769
-0.750 0.3487 0.01213 0.00404 -0.0823 0.7564 0.0847
-0.500 0.3760 0.01193 0.00385 -0.0819 0.7442 0.1014
-0.250 0.4028 0.01148 0.00371 -0.0817 0.7323 0.2059
0.000 0.4305 0.00942 0.00366 -0.0813 0.7208 1.0000
0.250 0.4576 0.00956 0.00357 -0.0809 0.7081 1.0000
0.500 0.4847 0.00969 0.00353 -0.0806 0.6955 1.0000
0.750 0.5118 0.00983 0.00352 -0.0803 0.6833 1.0000
1.000 0.5389 0.00998 0.00352 -0.0800 0.6713 1.0000
1.250 0.5660 0.01013 0.00353 -0.0797 0.6593 1.0000
1.500 0.5930 0.01028 0.00354 -0.0793 0.6477 1.0000
1.750 0.6201 0.01041 0.00360 -0.0791 0.6351 1.0000
2.000 0.6473 0.01056 0.00368 -0.0789 0.6231 1.0000
2.250 0.6744 0.01071 0.00376 -0.0787 0.6114 1.0000
2.500 0.7013 0.01087 0.00383 -0.0784 0.5992 1.0000
2.750 0.7283 0.01103 0.00390 -0.0781 0.5872 1.0000
3.000 0.7552 0.01118 0.00402 -0.0779 0.5751 1.0000
3.250 0.7822 0.01134 0.00416 -0.0777 0.5628 1.0000
3.500 0.8091 0.01151 0.00430 -0.0775 0.5505 1.0000
3.750 0.8358 0.01168 0.00445 -0.0772 0.5376 1.0000
4.000 0.8624 0.01187 0.00460 -0.0770 0.5256 1.0000
4.250 0.8890 0.01207 0.00476 -0.0767 0.5134 1.0000
4.500 0.9154 0.01227 0.00493 -0.0764 0.5012 1.0000
4.750 0.9418 0.01248 0.00517 -0.0762 0.4887 1.0000
5.000 0.9681 0.01270 0.00540 -0.0759 0.4762 1.0000
5.250 0.9941 0.01293 0.00564 -0.0756 0.4630 1.0000
5.500 1.0199 0.01318 0.00591 -0.0752 0.4497 1.0000
5.750 1.0456 0.01345 0.00619 -0.0749 0.4366 1.0000
6.000 1.0707 0.01371 0.00646 -0.0744 0.4207 1.0000
6.250 1.0947 0.01400 0.00669 -0.0738 0.4000 1.0000
6.500 1.1190 0.01428 0.00702 -0.0733 0.3783 1.0000
6.750 1.1427 0.01463 0.00735 -0.0727 0.3578 1.0000
7.000 1.1656 0.01502 0.00771 -0.0720 0.3298 1.0000
7.250 1.1882 0.01547 0.00812 -0.0713 0.3018 1.0000
7.500 1.2103 0.01599 0.00862 -0.0705 0.2724 1.0000
7.750 1.2307 0.01668 0.00916 -0.0696 0.2343 1.0000
8.000 1.2479 0.01769 0.00991 -0.0684 0.1931 1.0000
8.250 1.2644 0.01879 0.01083 -0.0670 0.1621 1.0000
8.500 1.2792 0.02007 0.01179 -0.0655 0.0936 1.0000
8.750 1.2839 0.02231 0.01355 -0.0627 0.0418 1.0000
9.000 1.2966 0.02366 0.01489 -0.0608 0.0327 1.0000
9.250 1.3080 0.02503 0.01636 -0.0586 0.0293 1.0000
9.500 1.3197 0.02624 0.01774 -0.0566 0.0279 1.0000
9.750 1.3267 0.02756 0.01923 -0.0539 0.0268 1.0000
10.000 1.3302 0.02907 0.02088 -0.0510 0.0260 1.0000
10.250 1.3321 0.03082 0.02275 -0.0484 0.0253 1.0000
10.500 1.3325 0.03282 0.02486 -0.0461 0.0248 1.0000
10.750 1.3319 0.03508 0.02723 -0.0441 0.0241 1.0000
11.000 1.3298 0.03766 0.02990 -0.0423 0.0234 1.0000
11.250 1.3264 0.04057 0.03288 -0.0405 0.0227 1.0000
11.500 1.3292 0.04303 0.03542 -0.0389 0.0222 1.0000
11.750 1.3353 0.04520 0.03772 -0.0376 0.0219 1.0000
12.000 1.3425 0.04739 0.04002 -0.0362 0.0217 1.0000
12.250 1.3508 0.04961 0.04236 -0.0347 0.0216 1.0000
12.500 1.3595 0.05194 0.04484 -0.0332 0.0215 1.0000
12.750 1.3676 0.05450 0.04757 -0.0318 0.0214 1.0000
13.000 1.3735 0.05741 0.05070 -0.0305 0.0215 1.0000
13.250 1.3766 0.06071 0.05423 -0.0294 0.0216 1.0000
13.500 1.3765 0.06447 0.05822 -0.0285 0.0218 1.0000
13.750 1.3732 0.06871 0.06269 -0.0279 0.0220 1.0000
14.000 1.3667 0.07333 0.06754 -0.0275 0.0222 1.0000
14.250 1.3605 0.07836 0.07280 -0.0273 0.0224 1.0000
14.500 1.3480 0.08295 0.07760 -0.0281 0.0225 1.0000
14.750 1.3342 0.08779 0.08265 -0.0294 0.0225 1.0000
15.000 1.3213 0.09259 0.08764 -0.0311 0.0226 1.0000
15.250 1.3058 0.09801 0.09325 -0.0334 0.0227 1.0000
15.500 1.2911 0.10354 0.09897 -0.0361 0.0228 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 137 (MVA H.15) AIRFOIL (goe137-il)