DAE-51 AIRFOIL (dae51-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: DAE-51 AIRFOIL (dae51-il) Reynolds number: 100,000 Max Cl/Cd: 63.29 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-dae51-il-100000-n5.txt Download as CSV file: xf-dae51-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.3448 0.09744 0.09273 -0.0404 1.0000 0.0544
-8.000 -0.3588 0.09602 0.09144 -0.0392 1.0000 0.0546
-7.750 -0.3755 0.09484 0.09038 -0.0371 1.0000 0.0547
-7.500 -0.3619 0.09005 0.08559 -0.0453 0.9926 0.0549
-7.250 -0.3387 0.08400 0.07948 -0.0556 0.9847 0.0550
-6.750 -0.3025 0.06951 0.06487 -0.0634 0.9735 0.0331
-6.500 -0.2776 0.06343 0.05868 -0.0717 0.9663 0.0330
-6.250 -0.2498 0.05695 0.05200 -0.0802 0.9597 0.0329
-6.000 -0.2240 0.05150 0.04635 -0.0862 0.9530 0.0325
-5.750 -0.1970 0.04632 0.04091 -0.0914 0.9463 0.0317
-5.500 -0.1660 0.04084 0.03502 -0.0968 0.9410 0.0310
-5.250 -0.1388 0.03627 0.02996 -0.1000 0.9335 0.0305
-5.000 -0.1055 0.03215 0.02524 -0.1032 0.9290 0.0304
-4.750 -0.0766 0.02909 0.02164 -0.1046 0.9224 0.0307
-4.500 -0.0451 0.02662 0.01859 -0.1059 0.9169 0.0318
-4.250 -0.0109 0.02464 0.01606 -0.1073 0.9130 0.0341
-4.000 0.0150 0.02325 0.01451 -0.1074 0.9055 0.0357
-3.750 0.0462 0.02192 0.01295 -0.1080 0.9004 0.0373
-3.500 0.0754 0.02081 0.01164 -0.1082 0.8946 0.0394
-3.250 0.1039 0.02007 0.01071 -0.1081 0.8880 0.0434
-3.000 0.1340 0.01904 0.00970 -0.1086 0.8834 0.0480
-2.750 0.1601 0.01842 0.00899 -0.1082 0.8757 0.0530
-2.500 0.1894 0.01777 0.00831 -0.1084 0.8701 0.0632
-2.250 0.2168 0.01723 0.00774 -0.1083 0.8632 0.0785
-2.000 0.2450 0.01657 0.00725 -0.1084 0.8567 0.1176
-1.750 0.2726 0.01572 0.00696 -0.1087 0.8509 0.2602
-1.500 0.2979 0.01504 0.00693 -0.1084 0.8434 0.4545
-1.250 0.3218 0.01437 0.00686 -0.1069 0.8382 0.6496
-1.000 0.3386 0.01381 0.00682 -0.1034 0.8297 0.8637
-0.750 0.3746 0.01363 0.00647 -0.1046 0.8240 1.0000
-0.500 0.4014 0.01374 0.00640 -0.1044 0.8154 1.0000
-0.250 0.4303 0.01376 0.00622 -0.1044 0.8093 1.0000
0.000 0.4570 0.01389 0.00621 -0.1042 0.8004 1.0000
0.250 0.4857 0.01390 0.00607 -0.1040 0.7942 1.0000
0.500 0.5122 0.01404 0.00611 -0.1038 0.7849 1.0000
0.750 0.5406 0.01406 0.00601 -0.1036 0.7782 1.0000
1.000 0.5671 0.01420 0.00608 -0.1033 0.7687 1.0000
1.250 0.5946 0.01427 0.00608 -0.1031 0.7606 1.0000
1.500 0.6220 0.01435 0.00609 -0.1028 0.7518 1.0000
1.750 0.6488 0.01447 0.00616 -0.1025 0.7422 1.0000
2.000 0.6767 0.01450 0.00612 -0.1022 0.7342 1.0000
2.250 0.7031 0.01463 0.00626 -0.1019 0.7234 1.0000
2.500 0.7299 0.01475 0.00635 -0.1016 0.7131 1.0000
2.750 0.7573 0.01481 0.00638 -0.1012 0.7035 1.0000
3.000 0.7840 0.01492 0.00650 -0.1009 0.6924 1.0000
3.250 0.8104 0.01505 0.00665 -0.1005 0.6804 1.0000
3.500 0.8369 0.01518 0.00678 -0.1000 0.6683 1.0000
3.750 0.8635 0.01530 0.00691 -0.0996 0.6558 1.0000
4.000 0.8899 0.01543 0.00707 -0.0992 0.6428 1.0000
4.250 0.9161 0.01558 0.00723 -0.0987 0.6289 1.0000
4.500 0.9421 0.01574 0.00742 -0.0982 0.6142 1.0000
4.750 0.9678 0.01593 0.00763 -0.0976 0.5985 1.0000
5.000 0.9932 0.01614 0.00790 -0.0970 0.5819 1.0000
5.250 1.0184 0.01637 0.00815 -0.0964 0.5644 1.0000
5.500 1.0432 0.01662 0.00841 -0.0957 0.5459 1.0000
5.750 1.0674 0.01692 0.00873 -0.0949 0.5257 1.0000
6.000 1.0911 0.01724 0.00910 -0.0941 0.5034 1.0000
6.250 1.1140 0.01762 0.00949 -0.0931 0.4786 1.0000
6.500 1.1359 0.01805 0.00990 -0.0921 0.4508 1.0000
6.750 1.1566 0.01856 0.01038 -0.0908 0.4187 1.0000
7.000 1.1758 0.01919 0.01090 -0.0894 0.3830 1.0000
7.250 1.1938 0.01991 0.01152 -0.0880 0.3439 1.0000
7.500 1.2102 0.02078 0.01224 -0.0864 0.3033 1.0000
7.750 1.2248 0.02179 0.01308 -0.0846 0.2613 1.0000
8.000 1.2378 0.02294 0.01406 -0.0828 0.2184 1.0000
8.250 1.2494 0.02422 0.01513 -0.0808 0.1775 1.0000
8.500 1.2596 0.02562 0.01632 -0.0788 0.1407 1.0000
8.750 1.2690 0.02705 0.01761 -0.0766 0.1101 1.0000
9.000 1.2760 0.02857 0.01900 -0.0742 0.0884 1.0000
9.250 1.2821 0.03007 0.02047 -0.0716 0.0741 1.0000
9.500 1.2871 0.03172 0.02213 -0.0691 0.0642 1.0000
9.750 1.2935 0.03330 0.02381 -0.0668 0.0563 1.0000
10.000 1.2969 0.03516 0.02570 -0.0646 0.0510 1.0000
10.250 1.3019 0.03695 0.02764 -0.0625 0.0468 1.0000
10.500 1.3044 0.03900 0.02975 -0.0606 0.0437 1.0000
10.750 1.3072 0.04112 0.03198 -0.0587 0.0409 1.0000
11.000 1.3122 0.04310 0.03412 -0.0571 0.0379 1.0000
11.250 1.3157 0.04526 0.03640 -0.0556 0.0357 1.0000
11.500 1.3174 0.04768 0.03889 -0.0541 0.0342 1.0000
11.750 1.3208 0.05009 0.04140 -0.0525 0.0328 1.0000
12.000 1.3266 0.05233 0.04386 -0.0512 0.0312 1.0000
12.250 1.3310 0.05471 0.04642 -0.0501 0.0295 1.0000
12.500 1.3339 0.05722 0.04906 -0.0491 0.0281 1.0000
12.750 1.3359 0.05986 0.05181 -0.0482 0.0270 1.0000
13.000 1.3385 0.06265 0.05467 -0.0472 0.0262 1.0000
13.250 1.3413 0.06567 0.05785 -0.0461 0.0255 1.0000
13.500 1.3416 0.06895 0.06142 -0.0454 0.0249 1.0000
13.750 1.3388 0.07261 0.06539 -0.0449 0.0244 1.0000
14.000 1.3333 0.07663 0.06968 -0.0448 0.0238 1.0000
14.250 1.3253 0.08097 0.07429 -0.0452 0.0233 1.0000
14.500 1.3155 0.08566 0.07923 -0.0459 0.0228 1.0000
14.750 1.3041 0.09077 0.08458 -0.0472 0.0224 1.0000
15.000 1.2913 0.09632 0.09037 -0.0491 0.0221 1.0000
15.250 1.2772 0.10240 0.09668 -0.0516 0.0219 1.0000
15.500 1.2614 0.10916 0.10366 -0.0548 0.0217 1.0000
15.750 1.2434 0.11679 0.11152 -0.0589 0.0217 1.0000
16.000 1.2222 0.12572 0.12070 -0.0642 0.0218 1.0000
16.250 1.1948 0.13711 0.13233 -0.0716 0.0222 1.0000
16.500 1.1569 0.15310 0.14853 -0.0824 0.0233 1.0000
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Polar data table (+)
Polar graphs
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