GOE 164 (MVA MK.10) AIRFOIL (goe164-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 164 (MVA MK.10) AIRFOIL (goe164-il) Reynolds number: 100,000 Max Cl/Cd: 62.58 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe164-il-100000-n5.txt Download as CSV file: xf-goe164-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 164 (MVA MK.10) 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.2691 0.10694 0.10253 -0.0261 1.0000 0.0462
-7.500 -0.2817 0.10634 0.10205 -0.0246 1.0000 0.0468
-7.250 -0.2886 0.10524 0.10105 -0.0253 0.9978 0.0472
-7.000 -0.2602 0.10067 0.09647 -0.0353 0.9897 0.0476
-6.750 -0.2334 0.09585 0.09164 -0.0431 0.9822 0.0477
-6.250 -0.1971 0.08455 0.08031 -0.0462 0.9711 0.0338
-6.000 -0.1723 0.08001 0.07576 -0.0525 0.9628 0.0335
-5.500 -0.1200 0.07076 0.06645 -0.0654 0.9437 0.0323
-5.250 -0.0919 0.06577 0.06143 -0.0723 0.9334 0.0310
-5.000 -0.0533 0.05962 0.05519 -0.0823 0.9260 0.0299
-4.750 -0.0166 0.05328 0.04875 -0.0918 0.9146 0.0294
-4.500 0.0305 0.04632 0.04160 -0.1032 0.9067 0.0303
-4.250 0.1129 0.03041 0.02462 -0.1278 0.9026 0.0324
-4.000 0.1641 0.02524 0.01840 -0.1352 0.8982 0.0344
-3.750 0.2007 0.02295 0.01565 -0.1376 0.8892 0.0378
-3.500 0.2408 0.02137 0.01378 -0.1401 0.8826 0.0431
-3.250 0.2768 0.02003 0.01206 -0.1416 0.8717 0.0485
-3.000 0.3155 0.01876 0.01061 -0.1437 0.8616 0.0581
-2.750 0.3562 0.01767 0.00935 -0.1460 0.8523 0.0725
-2.500 0.3951 0.01692 0.00866 -0.1479 0.8411 0.0986
-2.250 0.4323 0.01682 0.00847 -0.1492 0.8284 0.1428
-2.000 0.4690 0.01700 0.00838 -0.1503 0.8143 0.1711
-1.500 0.5371 0.01712 0.00810 -0.1521 0.7859 0.2061
-1.250 0.5688 0.01711 0.00793 -0.1526 0.7727 0.2198
-1.000 0.5997 0.01695 0.00767 -0.1529 0.7596 0.2262
-0.750 0.6301 0.01679 0.00731 -0.1532 0.7460 0.2310
-0.500 0.6592 0.01667 0.00706 -0.1532 0.7320 0.2359
-0.250 0.6874 0.01658 0.00687 -0.1530 0.7175 0.2408
0.250 0.7418 0.01646 0.00653 -0.1522 0.6868 0.2501
0.500 0.7679 0.01643 0.00645 -0.1516 0.6707 0.2554
0.750 0.7942 0.01645 0.00636 -0.1511 0.6545 0.2626
1.000 0.8202 0.01645 0.00634 -0.1504 0.6382 0.2696
1.250 0.8460 0.01650 0.00631 -0.1498 0.6219 0.2777
1.500 0.8716 0.01655 0.00632 -0.1492 0.6050 0.2857
1.750 0.8969 0.01663 0.00635 -0.1484 0.5876 0.2964
2.000 0.9219 0.01674 0.00643 -0.1477 0.5699 0.3110
2.250 0.9467 0.01686 0.00652 -0.1470 0.5523 0.3301
2.500 0.9713 0.01700 0.00664 -0.1462 0.5351 0.3580
2.750 0.9958 0.01709 0.00680 -0.1455 0.5183 0.4038
3.250 1.0397 0.01669 0.00710 -0.1429 0.4858 1.0000
3.500 1.0636 0.01702 0.00733 -0.1420 0.4691 1.0000
3.750 1.0870 0.01737 0.00756 -0.1411 0.4521 1.0000
4.000 1.1101 0.01774 0.00782 -0.1402 0.4353 1.0000
4.250 1.1332 0.01812 0.00813 -0.1393 0.4202 1.0000
4.500 1.1563 0.01851 0.00845 -0.1384 0.4074 1.0000
4.750 1.1793 0.01891 0.00879 -0.1375 0.3960 1.0000
5.000 1.2020 0.01934 0.00916 -0.1366 0.3851 1.0000
5.250 1.2250 0.01975 0.00960 -0.1357 0.3752 1.0000
5.500 1.2476 0.02020 0.01002 -0.1348 0.3668 1.0000
5.750 1.2702 0.02065 0.01049 -0.1339 0.3579 1.0000
6.000 1.2925 0.02112 0.01100 -0.1330 0.3498 1.0000
6.250 1.3146 0.02159 0.01152 -0.1320 0.3418 1.0000
6.500 1.3367 0.02209 0.01205 -0.1310 0.3350 1.0000
6.750 1.3587 0.02258 0.01263 -0.1301 0.3279 1.0000
7.000 1.3799 0.02313 0.01321 -0.1290 0.3215 1.0000
7.250 1.4015 0.02362 0.01389 -0.1280 0.3141 1.0000
7.500 1.4218 0.02419 0.01452 -0.1268 0.3073 1.0000
7.750 1.4425 0.02472 0.01521 -0.1257 0.2998 1.0000
8.000 1.4620 0.02530 0.01590 -0.1244 0.2928 1.0000
8.250 1.4816 0.02587 0.01664 -0.1231 0.2854 1.0000
8.500 1.4988 0.02648 0.01739 -0.1214 0.2766 1.0000
8.750 1.5136 0.02705 0.01811 -0.1194 0.2629 1.0000
9.000 1.5263 0.02767 0.01884 -0.1171 0.2456 1.0000
9.250 1.5358 0.02838 0.01960 -0.1143 0.2246 1.0000
9.500 1.5421 0.02935 0.02050 -0.1113 0.2023 1.0000
9.750 1.5470 0.03054 0.02164 -0.1082 0.1858 1.0000
10.000 1.5496 0.03200 0.02305 -0.1050 0.1702 1.0000
10.250 1.5497 0.03374 0.02475 -0.1018 0.1514 1.0000
10.500 1.5528 0.03542 0.02650 -0.0991 0.0976 1.0000
10.750 1.5358 0.03876 0.02945 -0.0951 0.0803 1.0000
11.000 1.5316 0.04130 0.03213 -0.0924 0.0608 1.0000
11.250 1.5175 0.04491 0.03562 -0.0897 0.0352 1.0000
11.500 1.5071 0.04845 0.03919 -0.0878 0.0297 1.0000
11.750 1.4970 0.05220 0.04307 -0.0864 0.0269 1.0000
12.000 1.4856 0.05634 0.04738 -0.0855 0.0253 1.0000
12.250 1.4764 0.06045 0.05169 -0.0851 0.0243 1.0000
12.500 1.4656 0.06497 0.05644 -0.0851 0.0235 1.0000
12.750 1.4533 0.06991 0.06160 -0.0855 0.0229 1.0000
13.000 1.4396 0.07532 0.06722 -0.0863 0.0223 1.0000
13.250 1.4248 0.08114 0.07326 -0.0875 0.0219 1.0000
13.500 1.4094 0.08726 0.07958 -0.0891 0.0216 1.0000
13.750 1.3936 0.09367 0.08617 -0.0909 0.0212 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 164 (MVA MK.10) AIRFOIL (goe164-il)