OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 50,000 Max Cl/Cd: 28.29 at α=8.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-oaf117-il-50000.txt Download as CSV file: xf-oaf117-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: OAF117 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
------ -------- --------- --------- -------- -------- --------
-7.750 -0.4969 0.10103 0.09397 0.0239 1.0000 0.3279
-7.500 -0.4937 0.09837 0.09135 0.0234 1.0000 0.3409
-7.250 -0.5601 0.07210 0.06504 -0.0172 1.0000 0.1938
-7.000 -0.5491 0.06546 0.05829 -0.0221 1.0000 0.1913
-6.750 -0.5380 0.05727 0.04985 -0.0298 1.0000 0.1899
-6.500 -0.5210 0.04813 0.04014 -0.0390 1.0000 0.1906
-6.250 -0.4973 0.04401 0.03580 -0.0417 1.0000 0.1977
-6.000 -0.4694 0.03917 0.03039 -0.0460 1.0000 0.2070
-5.750 -0.4430 0.03718 0.02838 -0.0464 1.0000 0.2185
-5.500 -0.4143 0.03459 0.02555 -0.0480 1.0000 0.2324
-5.250 -0.3858 0.03266 0.02349 -0.0489 1.0000 0.2499
-5.000 -0.3565 0.03092 0.02158 -0.0498 1.0000 0.2716
-4.750 -0.3283 0.02963 0.02025 -0.0501 1.0000 0.2945
-4.500 -0.3017 0.02885 0.01958 -0.0496 1.0000 0.3180
-4.250 -0.2733 0.02783 0.01849 -0.0499 1.0000 0.3447
-4.000 -0.2478 0.02727 0.01807 -0.0490 1.0000 0.3688
-3.750 -0.2199 0.02640 0.01714 -0.0493 1.0000 0.3959
-3.500 -0.1957 0.02583 0.01672 -0.0484 1.0000 0.4192
-3.250 -0.1694 0.02506 0.01592 -0.0486 1.0000 0.4456
-3.000 -0.1475 0.02445 0.01542 -0.0477 1.0000 0.4681
-2.750 -0.1290 0.02400 0.01509 -0.0465 1.0000 0.4906
-2.500 -0.1107 0.02362 0.01472 -0.0458 1.0000 0.5149
-2.250 -0.0944 0.02340 0.01458 -0.0449 1.0000 0.5376
-2.000 -0.0779 0.02332 0.01458 -0.0441 1.0000 0.5615
-1.750 -0.0592 0.02330 0.01460 -0.0441 1.0000 0.5883
-1.500 -0.0419 0.02342 0.01481 -0.0437 1.0000 0.6152
-1.250 -0.0229 0.02361 0.01509 -0.0437 1.0000 0.6451
-1.000 0.0324 0.02338 0.01495 -0.0490 0.9820 0.6937
-0.750 0.0851 0.02306 0.01473 -0.0537 0.9616 0.7420
-0.500 0.1345 0.02264 0.01446 -0.0570 0.9430 0.7941
-0.250 0.1752 0.02211 0.01410 -0.0581 0.9247 0.8564
0.000 0.2228 0.02149 0.01365 -0.0608 0.9018 0.9680
0.250 0.2905 0.02147 0.01334 -0.0710 0.8773 1.0000
0.500 0.3371 0.02177 0.01337 -0.0751 0.8561 1.0000
0.750 0.3710 0.02228 0.01369 -0.0766 0.8339 1.0000
1.000 0.4025 0.02280 0.01405 -0.0772 0.8137 1.0000
1.250 0.4319 0.02334 0.01446 -0.0772 0.7945 1.0000
1.500 0.4601 0.02389 0.01492 -0.0768 0.7760 1.0000
1.750 0.4874 0.02448 0.01542 -0.0763 0.7579 1.0000
2.000 0.5141 0.02507 0.01593 -0.0756 0.7400 1.0000
2.250 0.5405 0.02568 0.01649 -0.0749 0.7223 1.0000
2.500 0.5666 0.02629 0.01705 -0.0740 0.7046 1.0000
2.750 0.5925 0.02692 0.01764 -0.0732 0.6871 1.0000
3.000 0.6183 0.02754 0.01823 -0.0722 0.6695 1.0000
3.250 0.6440 0.02817 0.01884 -0.0713 0.6519 1.0000
3.500 0.6696 0.02881 0.01947 -0.0703 0.6344 1.0000
3.750 0.6952 0.02945 0.02012 -0.0693 0.6170 1.0000
4.000 0.7207 0.03012 0.02078 -0.0683 0.5995 1.0000
4.250 0.7462 0.03078 0.02146 -0.0673 0.5821 1.0000
4.500 0.7716 0.03144 0.02214 -0.0662 0.5645 1.0000
4.750 0.7970 0.03210 0.02282 -0.0651 0.5469 1.0000
5.000 0.8224 0.03271 0.02345 -0.0639 0.5291 1.0000
5.250 0.8478 0.03333 0.02409 -0.0627 0.5112 1.0000
5.500 0.8732 0.03386 0.02463 -0.0613 0.4931 1.0000
5.750 0.8988 0.03437 0.02512 -0.0600 0.4750 1.0000
6.000 0.9220 0.03555 0.02647 -0.0594 0.4536 1.0000
6.250 0.9462 0.03631 0.02730 -0.0583 0.4332 1.0000
6.500 0.9712 0.03679 0.02778 -0.0569 0.4133 1.0000
6.750 0.9971 0.03695 0.02790 -0.0552 0.3935 1.0000
7.000 1.0222 0.03730 0.02819 -0.0537 0.3728 1.0000
7.250 1.0444 0.03822 0.02924 -0.0526 0.3496 1.0000
7.500 1.0702 0.03825 0.02913 -0.0509 0.3278 1.0000
7.750 1.0928 0.03902 0.02992 -0.0497 0.3048 1.0000
8.000 1.1155 0.03971 0.03058 -0.0483 0.2816 1.0000
8.250 1.1400 0.04030 0.03092 -0.0469 0.2601 1.0000
8.500 1.1572 0.04226 0.03312 -0.0460 0.2388 1.0000
8.750 1.1765 0.04395 0.03485 -0.0449 0.2204 1.0000
9.000 1.1947 0.04598 0.03694 -0.0438 0.2049 1.0000
9.250 1.2100 0.04858 0.03970 -0.0429 0.1921 1.0000
9.500 1.2275 0.05102 0.04217 -0.0420 0.1815 1.0000
9.750 1.2387 0.05409 0.04549 -0.0410 0.1724 1.0000
10.000 1.2424 0.05823 0.04994 -0.0402 0.1663 1.0000
10.250 1.2268 0.06416 0.05641 -0.0396 0.1633 1.0000
10.500 1.1979 0.07112 0.06377 -0.0397 0.1626 1.0000
10.750 1.1453 0.08015 0.07307 -0.0415 0.1654 1.0000
11.000 1.0956 0.09166 0.08465 -0.0475 0.1682 1.0000
|
Polar data table (+)
Polar graphs
<< Back to OAF117 AIRFOIL (oaf117-il)