EPPLER 399 AIRFOIL (e399-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 399 AIRFOIL (e399-il) Reynolds number: 100,000 Max Cl/Cd: 54 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e399-il-100000-n5.txt Download as CSV file: xf-e399-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 399 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
------ -------- --------- --------- -------- -------- --------
-11.250 -0.4253 0.07788 0.07308 -0.0785 0.9880 0.0262
-11.000 -0.4452 0.06748 0.06251 -0.0886 0.9791 0.0259
-10.750 -0.4596 0.05881 0.05363 -0.0980 0.9677 0.0257
-10.500 -0.4729 0.05108 0.04557 -0.1072 0.9538 0.0261
-10.250 -0.4719 0.04613 0.04043 -0.1140 0.9420 0.0268
-10.000 -0.4638 0.04242 0.03656 -0.1197 0.9302 0.0279
-9.750 -0.4531 0.03900 0.03285 -0.1240 0.9192 0.0293
-9.500 -0.4316 0.03573 0.02922 -0.1284 0.9131 0.0314
-9.250 -0.4153 0.03397 0.02732 -0.1298 0.9029 0.0340
-9.000 -0.3883 0.03187 0.02500 -0.1328 0.8975 0.0377
-8.750 -0.3690 0.03026 0.02309 -0.1336 0.8882 0.0421
-8.500 -0.3403 0.02880 0.02152 -0.1358 0.8826 0.0472
-8.250 -0.3160 0.02760 0.02020 -0.1369 0.8750 0.0528
-8.000 -0.2890 0.02641 0.01886 -0.1382 0.8683 0.0595
-7.750 -0.2564 0.02522 0.01749 -0.1403 0.8639 0.0677
-7.500 -0.2343 0.02444 0.01655 -0.1402 0.8549 0.0755
-7.250 -0.2032 0.02355 0.01554 -0.1417 0.8496 0.0852
-7.000 -0.1763 0.02282 0.01474 -0.1423 0.8429 0.0950
-6.750 -0.1489 0.02214 0.01398 -0.1430 0.8361 0.1057
-6.500 -0.1164 0.02143 0.01316 -0.1444 0.8315 0.1186
-6.250 -0.0917 0.02096 0.01261 -0.1444 0.8238 0.1311
-6.000 -0.0620 0.02044 0.01201 -0.1452 0.8179 0.1461
-5.750 -0.0297 0.01995 0.01143 -0.1464 0.8134 0.1630
-5.500 -0.0055 0.01967 0.01110 -0.1462 0.8059 0.1785
-5.250 0.0249 0.01931 0.01067 -0.1470 0.8006 0.1968
-5.000 0.0551 0.01902 0.01029 -0.1476 0.7955 0.2158
-4.750 0.0808 0.01885 0.01007 -0.1475 0.7886 0.2340
-4.500 0.1112 0.01863 0.00979 -0.1482 0.7837 0.2533
-4.250 0.1408 0.01847 0.00957 -0.1487 0.7789 0.2721
-4.000 0.1663 0.01840 0.00947 -0.1484 0.7723 0.2899
-3.750 0.1961 0.01827 0.00928 -0.1489 0.7674 0.3085
-3.500 0.2264 0.01816 0.00909 -0.1494 0.7630 0.3268
-3.250 0.2510 0.01817 0.00908 -0.1490 0.7566 0.3434
-3.000 0.2798 0.01811 0.00896 -0.1492 0.7518 0.3607
-2.750 0.3111 0.01801 0.00878 -0.1498 0.7480 0.3783
-2.500 0.3352 0.01807 0.00885 -0.1493 0.7417 0.3941
-2.250 0.3626 0.01807 0.00882 -0.1493 0.7366 0.4106
-2.000 0.3928 0.01802 0.00871 -0.1497 0.7327 0.4277
-1.750 0.4194 0.01806 0.00874 -0.1496 0.7278 0.4440
-1.500 0.4447 0.01813 0.00882 -0.1492 0.7222 0.4603
-1.250 0.4732 0.01813 0.00880 -0.1493 0.7180 0.4773
-1.000 0.5038 0.01810 0.00873 -0.1498 0.7146 0.4949
-0.750 0.5262 0.01827 0.00895 -0.1490 0.7086 0.5114
-0.500 0.5527 0.01834 0.00904 -0.1487 0.7038 0.5289
-0.250 0.5818 0.01835 0.00905 -0.1489 0.7001 0.5475
0.000 0.6088 0.01843 0.00915 -0.1488 0.6959 0.5661
0.250 0.6316 0.01861 0.00941 -0.1480 0.6904 0.5849
0.500 0.6584 0.01868 0.00951 -0.1478 0.6860 0.6051
0.750 0.6879 0.01870 0.00953 -0.1480 0.6826 0.6270
1.000 0.7111 0.01891 0.00981 -0.1472 0.6776 0.6484
1.250 0.7341 0.01908 0.01007 -0.1463 0.6725 0.6703
1.500 0.7607 0.01914 0.01018 -0.1460 0.6685 0.6945
1.750 0.7895 0.01915 0.01021 -0.1459 0.6654 0.7203
2.000 0.8071 0.01945 0.01065 -0.1442 0.6595 0.7463
2.250 0.8285 0.01958 0.01087 -0.1429 0.6547 0.7753
2.500 0.8525 0.01959 0.01094 -0.1418 0.6510 0.8083
2.750 0.8734 0.01962 0.01106 -0.1402 0.6470 0.8483
3.000 0.8859 0.01981 0.01141 -0.1373 0.6411 0.9095
3.250 0.9163 0.01989 0.01150 -0.1380 0.6366 1.0000
3.500 0.9499 0.02000 0.01154 -0.1392 0.6331 1.0000
3.750 0.9711 0.02049 0.01207 -0.1386 0.6272 1.0000
4.000 0.9963 0.02081 0.01240 -0.1384 0.6219 1.0000
4.250 1.0274 0.02095 0.01252 -0.1391 0.6179 1.0000
4.500 1.0520 0.02130 0.01290 -0.1388 0.6128 1.0000
4.750 1.0728 0.02175 0.01341 -0.1380 0.6066 1.0000
5.000 1.1025 0.02189 0.01356 -0.1384 0.6021 1.0000
5.250 1.1278 0.02219 0.01390 -0.1381 0.5969 1.0000
5.500 1.1464 0.02267 0.01447 -0.1368 0.5901 1.0000
5.750 1.1767 0.02275 0.01457 -0.1373 0.5853 1.0000
6.000 1.1959 0.02320 0.01511 -0.1360 0.5786 1.0000
6.250 1.2175 0.02352 0.01550 -0.1351 0.5718 1.0000
6.500 1.2511 0.02344 0.01543 -0.1359 0.5668 1.0000
6.750 1.2597 0.02415 0.01630 -0.1331 0.5578 1.0000
7.000 1.2914 0.02406 0.01623 -0.1336 0.5517 1.0000
7.250 1.3000 0.02472 0.01705 -0.1307 0.5425 1.0000
7.500 1.3301 0.02463 0.01698 -0.1309 0.5355 1.0000
7.750 1.3351 0.02533 0.01783 -0.1274 0.5257 1.0000
8.000 1.3605 0.02534 0.01788 -0.1268 0.5175 1.0000
8.250 1.3655 0.02599 0.01867 -0.1234 0.5072 1.0000
8.500 1.3769 0.02648 0.01924 -0.1209 0.4969 1.0000
8.750 1.3970 0.02661 0.01941 -0.1195 0.4866 1.0000
9.000 1.3997 0.02750 0.02042 -0.1161 0.4743 1.0000
9.250 1.4061 0.02830 0.02133 -0.1133 0.4616 1.0000
9.500 1.4144 0.02907 0.02216 -0.1109 0.4482 1.0000
9.750 1.4220 0.02993 0.02308 -0.1084 0.4338 1.0000
10.000 1.4279 0.03095 0.02415 -0.1060 0.4183 1.0000
10.250 1.4325 0.03216 0.02539 -0.1035 0.4016 1.0000
10.500 1.4362 0.03350 0.02676 -0.1012 0.3836 1.0000
10.750 1.4393 0.03498 0.02823 -0.0989 0.3644 1.0000
11.000 1.4414 0.03662 0.02983 -0.0968 0.3445 1.0000
11.250 1.4405 0.03864 0.03184 -0.0947 0.3238 1.0000
11.500 1.4390 0.04080 0.03395 -0.0927 0.3028 1.0000
11.750 1.4365 0.04315 0.03623 -0.0908 0.2822 1.0000
12.000 1.4331 0.04575 0.03880 -0.0892 0.2617 1.0000
12.250 1.4290 0.04853 0.04154 -0.0877 0.2419 1.0000
12.500 1.4243 0.05149 0.04444 -0.0864 0.2233 1.0000
12.750 1.4197 0.05459 0.04750 -0.0853 0.2058 1.0000
13.000 1.4156 0.05778 0.05067 -0.0845 0.1887 1.0000
13.250 1.4116 0.06108 0.05395 -0.0838 0.1728 1.0000
13.500 1.4077 0.06450 0.05736 -0.0834 0.1578 1.0000
13.750 1.4038 0.06802 0.06088 -0.0830 0.1438 1.0000
14.000 1.3998 0.07165 0.06451 -0.0829 0.1311 1.0000
14.250 1.3956 0.07543 0.06829 -0.0828 0.1195 1.0000
14.500 1.3924 0.07918 0.07208 -0.0830 0.1085 1.0000
14.750 1.3899 0.08292 0.07587 -0.0832 0.0985 1.0000
15.000 1.3868 0.08681 0.07980 -0.0836 0.0898 1.0000
15.250 1.3821 0.09100 0.08399 -0.0841 0.0825 1.0000
15.500 1.3802 0.09490 0.08798 -0.0847 0.0750 1.0000
15.750 1.3772 0.09899 0.09213 -0.0855 0.0689 1.0000
16.000 1.3732 0.10329 0.09648 -0.0865 0.0635 1.0000
16.250 1.3713 0.10732 0.10061 -0.0874 0.0583 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 399 AIRFOIL (e399-il)