DAE-51 AIRFOIL (dae51-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: DAE-51 AIRFOIL (dae51-il) Reynolds number: 50,000 Max Cl/Cd: 40.97 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae51-il-50000-n5.txt Download as CSV file: xf-dae51-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: DAE-51 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
------ -------- --------- --------- -------- -------- --------
-8.000 -0.3345 0.10032 0.09367 -0.0346 1.0000 0.1176
-7.750 -0.3386 0.09820 0.09167 -0.0342 1.0000 0.1198
-7.500 -0.3517 0.09692 0.09054 -0.0328 1.0000 0.1222
-7.250 -0.3824 0.09723 0.09109 -0.0346 1.0000 0.1251
-7.000 -0.3926 0.09510 0.08907 -0.0352 1.0000 0.1257
-6.750 -0.3903 0.09177 0.08578 -0.0320 1.0000 0.1266
-6.250 -0.3878 0.08104 0.07498 -0.0358 1.0000 0.0654
-5.750 -0.3722 0.07010 0.06381 -0.0465 0.9978 0.0544
-5.500 -0.3473 0.06556 0.05916 -0.0507 0.9920 0.0531
-5.250 -0.3167 0.06011 0.05348 -0.0575 0.9858 0.0523
-5.000 -0.2834 0.05455 0.04757 -0.0646 0.9795 0.0525
-4.750 -0.2465 0.04923 0.04174 -0.0713 0.9738 0.0531
-4.500 -0.2096 0.04463 0.03659 -0.0765 0.9686 0.0531
-4.250 -0.1737 0.04073 0.03213 -0.0805 0.9629 0.0530
-4.000 -0.1337 0.03725 0.02797 -0.0844 0.9585 0.0534
-3.750 -0.0970 0.03457 0.02461 -0.0869 0.9531 0.0554
-3.500 -0.0633 0.03275 0.02264 -0.0891 0.9474 0.0595
-3.250 -0.0234 0.03085 0.02028 -0.0917 0.9432 0.0629
-3.000 0.0093 0.02938 0.01837 -0.0926 0.9368 0.0665
-2.750 0.0444 0.02819 0.01706 -0.0942 0.9313 0.0741
-2.500 0.0825 0.02709 0.01567 -0.0959 0.9268 0.0827
-2.250 0.1128 0.02635 0.01483 -0.0965 0.9193 0.0971
-2.000 0.1525 0.02543 0.01387 -0.0989 0.9145 0.1249
-1.750 0.1868 0.02433 0.01322 -0.1008 0.9081 0.2022
-1.500 0.2183 0.02275 0.01312 -0.1019 0.9024 0.5304
-1.250 0.2380 0.02162 0.01291 -0.0982 0.8949 1.0000
-1.000 0.2704 0.02184 0.01268 -0.0993 0.8867 1.0000
-0.750 0.3026 0.02208 0.01255 -0.1003 0.8790 1.0000
-0.500 0.3347 0.02231 0.01246 -0.1012 0.8710 1.0000
-0.250 0.3640 0.02258 0.01248 -0.1017 0.8624 1.0000
0.000 0.3972 0.02277 0.01245 -0.1027 0.8550 1.0000
0.250 0.4239 0.02309 0.01258 -0.1027 0.8454 1.0000
0.500 0.4583 0.02323 0.01254 -0.1037 0.8388 1.0000
0.750 0.4829 0.02358 0.01277 -0.1033 0.8283 1.0000
1.000 0.5131 0.02380 0.01287 -0.1037 0.8202 1.0000
1.250 0.5416 0.02404 0.01302 -0.1038 0.8112 1.0000
1.500 0.5674 0.02436 0.01328 -0.1035 0.8011 1.0000
1.750 0.6001 0.02443 0.01329 -0.1040 0.7941 1.0000
2.000 0.6237 0.02481 0.01364 -0.1033 0.7828 1.0000
2.250 0.6501 0.02509 0.01389 -0.1030 0.7728 1.0000
2.500 0.6812 0.02514 0.01395 -0.1031 0.7649 1.0000
2.750 0.7048 0.02551 0.01432 -0.1024 0.7533 1.0000
3.000 0.7305 0.02577 0.01461 -0.1019 0.7426 1.0000
3.250 0.7622 0.02570 0.01456 -0.1019 0.7347 1.0000
3.500 0.7857 0.02604 0.01498 -0.1011 0.7224 1.0000
3.750 0.8101 0.02632 0.01532 -0.1003 0.7104 1.0000
4.000 0.8360 0.02650 0.01557 -0.0997 0.6988 1.0000
4.250 0.8642 0.02651 0.01569 -0.0991 0.6882 1.0000
4.500 0.8919 0.02652 0.01579 -0.0984 0.6768 1.0000
4.750 0.9164 0.02671 0.01609 -0.0975 0.6633 1.0000
5.000 0.9413 0.02686 0.01636 -0.0966 0.6495 1.0000
5.250 0.9666 0.02696 0.01662 -0.0956 0.6353 1.0000
5.500 0.9920 0.02704 0.01682 -0.0946 0.6206 1.0000
5.750 1.0177 0.02708 0.01699 -0.0936 0.6052 1.0000
6.000 1.0437 0.02707 0.01711 -0.0925 0.5890 1.0000
6.250 1.0664 0.02729 0.01750 -0.0912 0.5703 1.0000
6.500 1.0891 0.02749 0.01784 -0.0898 0.5503 1.0000
6.750 1.1141 0.02750 0.01795 -0.0884 0.5298 1.0000
7.000 1.1336 0.02788 0.01848 -0.0867 0.5054 1.0000
7.250 1.1536 0.02819 0.01891 -0.0849 0.4793 1.0000
7.500 1.1717 0.02860 0.01938 -0.0829 0.4500 1.0000
7.750 1.1874 0.02914 0.01994 -0.0807 0.4173 1.0000
8.000 1.2005 0.02984 0.02059 -0.0783 0.3817 1.0000
8.250 1.2099 0.03082 0.02155 -0.0756 0.3432 1.0000
8.500 1.2168 0.03205 0.02266 -0.0729 0.3042 1.0000
8.750 1.2203 0.03353 0.02397 -0.0700 0.2668 1.0000
9.000 1.2212 0.03526 0.02558 -0.0671 0.2305 1.0000
9.250 1.2206 0.03734 0.02748 -0.0644 0.1972 1.0000
9.500 1.2193 0.03968 0.02965 -0.0621 0.1681 1.0000
9.750 1.2184 0.04217 0.03205 -0.0601 0.1432 1.0000
10.000 1.2175 0.04481 0.03458 -0.0582 0.1242 1.0000
10.250 1.2177 0.04746 0.03724 -0.0566 0.1091 1.0000
10.500 1.2183 0.05015 0.03991 -0.0550 0.0974 1.0000
10.750 1.2193 0.05285 0.04259 -0.0536 0.0885 1.0000
11.000 1.2238 0.05537 0.04525 -0.0522 0.0799 1.0000
11.250 1.2281 0.05792 0.04783 -0.0509 0.0740 1.0000
11.500 1.2368 0.06024 0.05038 -0.0494 0.0683 1.0000
11.750 1.2434 0.06267 0.05272 -0.0481 0.0637 1.0000
12.000 1.2515 0.06530 0.05575 -0.0469 0.0591 1.0000
12.250 1.2598 0.06792 0.05860 -0.0457 0.0559 1.0000
12.500 1.2703 0.07042 0.06112 -0.0445 0.0532 1.0000
12.750 1.2729 0.07395 0.06497 -0.0437 0.0512 1.0000
13.000 1.2687 0.07802 0.06942 -0.0434 0.0493 1.0000
13.250 1.2620 0.08237 0.07408 -0.0434 0.0478 1.0000
13.500 1.2534 0.08700 0.07898 -0.0439 0.0466 1.0000
13.750 1.2432 0.09199 0.08421 -0.0449 0.0457 1.0000
14.000 1.2304 0.09755 0.09002 -0.0464 0.0452 1.0000
14.250 1.2141 0.10390 0.09662 -0.0487 0.0450 1.0000
14.500 1.1936 0.11143 0.10440 -0.0524 0.0451 1.0000
14.750 1.1674 0.12090 0.11412 -0.0579 0.0455 1.0000
15.000 1.1322 0.13402 0.12748 -0.0665 0.0468 1.0000
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Polar data table (+)
Polar graphs
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