OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: OAF117 AIRFOIL (oaf117-il) Reynolds number: 50,000 Max Cl/Cd: 34.84 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-oaf117-il-50000-n5.txt Download as CSV file: xf-oaf117-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.6808 0.07496 0.06763 -0.0158 1.0000 0.0803
-9.000 -0.7114 0.06119 0.05352 -0.0309 1.0000 0.0802
-8.750 -0.7189 0.05215 0.04383 -0.0389 1.0000 0.0813
-8.500 -0.6994 0.05027 0.04198 -0.0392 1.0000 0.0837
-8.250 -0.6822 0.04707 0.03854 -0.0411 1.0000 0.0870
-8.000 -0.6647 0.04238 0.03323 -0.0443 1.0000 0.0908
-7.750 -0.6423 0.03945 0.02999 -0.0457 1.0000 0.0943
-7.500 -0.6179 0.03771 0.02815 -0.0464 1.0000 0.0989
-7.250 -0.5922 0.03479 0.02463 -0.0481 1.0000 0.1048
-7.000 -0.5666 0.03352 0.02348 -0.0483 1.0000 0.1097
-6.750 -0.5391 0.03156 0.02110 -0.0493 1.0000 0.1176
-6.500 -0.5126 0.03033 0.01996 -0.0497 1.0000 0.1244
-6.250 -0.4847 0.02883 0.01825 -0.0503 1.0000 0.1342
-6.000 -0.4569 0.02767 0.01702 -0.0508 1.0000 0.1450
-5.750 -0.4290 0.02667 0.01604 -0.0513 1.0000 0.1572
-5.500 -0.4007 0.02570 0.01505 -0.0518 1.0000 0.1722
-5.250 -0.3723 0.02481 0.01412 -0.0523 1.0000 0.1894
-5.000 -0.3436 0.02404 0.01331 -0.0528 1.0000 0.2087
-4.750 -0.3155 0.02343 0.01280 -0.0531 1.0000 0.2280
-4.500 -0.2873 0.02292 0.01234 -0.0535 1.0000 0.2485
-4.250 -0.2592 0.02245 0.01190 -0.0538 1.0000 0.2704
-4.000 -0.2312 0.02202 0.01150 -0.0541 1.0000 0.2927
-3.750 -0.2039 0.02165 0.01121 -0.0542 1.0000 0.3131
-3.500 -0.1769 0.02130 0.01093 -0.0543 1.0000 0.3337
-3.250 -0.1509 0.02099 0.01067 -0.0543 1.0000 0.3542
-3.000 -0.1267 0.02075 0.01047 -0.0542 1.0000 0.3746
-2.750 -0.1049 0.02062 0.01039 -0.0538 0.9998 0.3942
-2.500 -0.0601 0.02039 0.01021 -0.0573 0.9756 0.4198
-2.250 -0.0154 0.02017 0.00998 -0.0606 0.9524 0.4481
-2.000 0.0254 0.02000 0.00986 -0.0629 0.9283 0.4756
-1.750 0.0624 0.01988 0.00975 -0.0643 0.9036 0.5038
-1.500 0.0948 0.01979 0.00966 -0.0646 0.8788 0.5297
-1.250 0.1243 0.01972 0.00958 -0.0642 0.8554 0.5543
-1.000 0.1511 0.01966 0.00954 -0.0634 0.8320 0.5792
-0.750 0.1770 0.01961 0.00952 -0.0622 0.8105 0.6071
-0.500 0.2022 0.01956 0.00949 -0.0609 0.7904 0.6381
-0.250 0.2262 0.01948 0.00948 -0.0593 0.7715 0.6714
0.000 0.2492 0.01938 0.00947 -0.0575 0.7532 0.7090
0.250 0.2709 0.01923 0.00943 -0.0553 0.7354 0.7540
0.500 0.2909 0.01900 0.00932 -0.0526 0.7179 0.8157
0.750 0.3176 0.01860 0.00901 -0.0514 0.6998 1.0000
1.000 0.3489 0.01885 0.00904 -0.0521 0.6825 1.0000
1.250 0.3792 0.01913 0.00913 -0.0526 0.6658 1.0000
1.500 0.4090 0.01943 0.00928 -0.0529 0.6498 1.0000
1.750 0.4385 0.01974 0.00947 -0.0531 0.6342 1.0000
2.000 0.4678 0.02006 0.00969 -0.0533 0.6190 1.0000
2.250 0.4968 0.02040 0.00994 -0.0533 0.6044 1.0000
2.500 0.5255 0.02074 0.01019 -0.0533 0.5898 1.0000
2.750 0.5539 0.02108 0.01044 -0.0531 0.5756 1.0000
3.000 0.5822 0.02143 0.01073 -0.0530 0.5611 1.0000
3.250 0.6106 0.02181 0.01108 -0.0529 0.5454 1.0000
3.500 0.6388 0.02219 0.01144 -0.0528 0.5299 1.0000
3.750 0.6668 0.02258 0.01183 -0.0527 0.5143 1.0000
4.000 0.6947 0.02298 0.01221 -0.0525 0.4988 1.0000
4.250 0.7224 0.02338 0.01261 -0.0523 0.4834 1.0000
4.500 0.7499 0.02380 0.01304 -0.0521 0.4682 1.0000
4.750 0.7773 0.02423 0.01347 -0.0518 0.4531 1.0000
5.000 0.8044 0.02466 0.01391 -0.0515 0.4378 1.0000
5.250 0.8314 0.02511 0.01438 -0.0512 0.4224 1.0000
5.500 0.8582 0.02557 0.01486 -0.0509 0.4067 1.0000
5.750 0.8847 0.02606 0.01538 -0.0505 0.3906 1.0000
6.000 0.9109 0.02659 0.01596 -0.0502 0.3740 1.0000
6.250 0.9367 0.02715 0.01658 -0.0498 0.3566 1.0000
6.500 0.9621 0.02774 0.01725 -0.0495 0.3386 1.0000
6.750 0.9869 0.02837 0.01797 -0.0491 0.3199 1.0000
7.000 1.0111 0.02903 0.01868 -0.0486 0.3007 1.0000
7.250 1.0347 0.02970 0.01937 -0.0481 0.2815 1.0000
7.500 1.0575 0.03044 0.02009 -0.0475 0.2625 1.0000
7.750 1.0793 0.03135 0.02104 -0.0470 0.2427 1.0000
8.000 1.1002 0.03236 0.02214 -0.0465 0.2225 1.0000
8.250 1.1200 0.03344 0.02323 -0.0459 0.2040 1.0000
8.500 1.1384 0.03468 0.02447 -0.0452 0.1869 1.0000
8.750 1.1555 0.03608 0.02587 -0.0444 0.1712 1.0000
9.000 1.1713 0.03766 0.02746 -0.0436 0.1568 1.0000
9.250 1.1859 0.03938 0.02925 -0.0427 0.1442 1.0000
9.500 1.1991 0.04121 0.03118 -0.0417 0.1333 1.0000
9.750 1.2113 0.04306 0.03301 -0.0407 0.1247 1.0000
10.000 1.2220 0.04508 0.03516 -0.0396 0.1165 1.0000
10.250 1.2317 0.04719 0.03736 -0.0385 0.1101 1.0000
10.500 1.2394 0.04941 0.03975 -0.0374 0.1043 1.0000
10.750 1.2470 0.05151 0.04182 -0.0362 0.0999 1.0000
11.000 1.2472 0.05431 0.04493 -0.0348 0.0958 1.0000
11.250 1.2483 0.05699 0.04776 -0.0339 0.0924 1.0000
11.500 1.2545 0.05923 0.04994 -0.0331 0.0894 1.0000
11.750 1.2486 0.06312 0.05413 -0.0331 0.0869 1.0000
12.000 1.2395 0.06761 0.05892 -0.0339 0.0848 1.0000
12.250 1.2298 0.07236 0.06391 -0.0352 0.0829 1.0000
12.500 1.2199 0.07734 0.06909 -0.0368 0.0814 1.0000
12.750 1.2112 0.08232 0.07420 -0.0385 0.0800 1.0000
13.000 1.2177 0.08488 0.07674 -0.0383 0.0780 1.0000
13.250 1.1886 0.09361 0.08575 -0.0431 0.0777 1.0000
13.500 1.1523 0.10449 0.09690 -0.0496 0.0777 1.0000
13.750 1.1041 0.11918 0.11184 -0.0589 0.0780 1.0000
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Polar data table (+)
Polar graphs
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