EPPLER 421 AIRFOIL (e421-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 421 AIRFOIL (e421-il) Reynolds number: 200,000 Max Cl/Cd: 74.79 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e421-il-200000-n5.txt Download as CSV file: xf-e421-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 421 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 0.1932 0.09829 0.09295 -0.1240 0.6836 0.0298
-9.750 0.1986 0.09441 0.08900 -0.1243 0.6743 0.0222
-9.250 0.2047 0.08805 0.08256 -0.1260 0.6575 0.0192
-8.750 0.2213 0.08405 0.07849 -0.1265 0.6412 0.0187
-8.500 0.2257 0.08138 0.07580 -0.1272 0.6342 0.0187
-8.250 0.2318 0.07912 0.07350 -0.1276 0.6276 0.0184
-8.000 0.2366 0.07660 0.07098 -0.1282 0.6213 0.0183
-7.750 0.2399 0.07393 0.06831 -0.1287 0.6155 0.0182
-7.250 0.2431 0.06835 0.06273 -0.1297 0.6051 0.0182
-7.000 0.2339 0.06432 0.05874 -0.1303 0.6007 0.0190
-6.500 0.1744 0.02633 0.01964 -0.1722 0.5990 0.0190
-6.250 0.1995 0.02432 0.01729 -0.1741 0.5934 0.0195
-6.000 0.2252 0.02310 0.01587 -0.1750 0.5879 0.0201
-5.750 0.2511 0.02193 0.01448 -0.1756 0.5820 0.0207
-5.500 0.2771 0.02089 0.01315 -0.1760 0.5765 0.0217
-5.250 0.3036 0.02002 0.01202 -0.1762 0.5719 0.0227
-5.000 0.3303 0.01939 0.01131 -0.1764 0.5669 0.0238
-4.750 0.3573 0.01879 0.01053 -0.1765 0.5621 0.0254
-4.500 0.3842 0.01832 0.00992 -0.1765 0.5578 0.0277
-4.000 0.4388 0.01749 0.00884 -0.1765 0.5497 0.0349
-3.750 0.4661 0.01712 0.00838 -0.1765 0.5453 0.0398
-3.500 0.4932 0.01685 0.00802 -0.1764 0.5411 0.0464
-3.250 0.5205 0.01665 0.00771 -0.1762 0.5374 0.0545
-3.000 0.5479 0.01649 0.00746 -0.1761 0.5341 0.0636
-2.750 0.5753 0.01635 0.00731 -0.1760 0.5305 0.0731
-2.500 0.6026 0.01625 0.00717 -0.1759 0.5270 0.0828
-2.250 0.6298 0.01619 0.00705 -0.1757 0.5237 0.0926
-2.000 0.6570 0.01618 0.00695 -0.1755 0.5205 0.1025
-1.750 0.6845 0.01619 0.00689 -0.1753 0.5176 0.1123
-1.500 0.7114 0.01617 0.00687 -0.1751 0.5144 0.1218
-1.250 0.7384 0.01617 0.00683 -0.1748 0.5110 0.1317
-1.000 0.7653 0.01621 0.00685 -0.1746 0.5080 0.1415
-0.750 0.7923 0.01624 0.00685 -0.1744 0.5053 0.1510
-0.500 0.8194 0.01632 0.00684 -0.1741 0.5027 0.1612
-0.250 0.8467 0.01640 0.00688 -0.1740 0.5003 0.1705
0.000 0.8736 0.01649 0.00691 -0.1737 0.4977 0.1803
0.250 0.8997 0.01653 0.00701 -0.1734 0.4947 0.1895
0.500 0.9261 0.01662 0.00706 -0.1730 0.4918 0.1992
0.750 0.9524 0.01669 0.00715 -0.1728 0.4892 0.2085
1.000 0.9790 0.01680 0.00721 -0.1725 0.4868 0.2185
1.250 1.0056 0.01689 0.00731 -0.1723 0.4846 0.2283
1.500 1.0327 0.01703 0.00738 -0.1721 0.4824 0.2388
1.750 1.0588 0.01714 0.00753 -0.1718 0.4801 0.2496
2.000 1.0841 0.01725 0.00769 -0.1714 0.4775 0.2609
2.250 1.1095 0.01737 0.00786 -0.1709 0.4749 0.2738
2.500 1.1350 0.01750 0.00803 -0.1706 0.4726 0.2875
2.750 1.1607 0.01763 0.00820 -0.1702 0.4704 0.3034
3.000 1.1863 0.01776 0.00837 -0.1699 0.4682 0.3230
3.250 1.2124 0.01791 0.00853 -0.1697 0.4660 0.3472
3.500 1.2392 0.01806 0.00871 -0.1696 0.4640 0.3809
3.750 1.2626 0.01815 0.00899 -0.1689 0.4617 0.4369
4.250 1.3096 0.01783 0.00951 -0.1673 0.4572 1.0000
4.500 1.3337 0.01809 0.00976 -0.1668 0.4549 1.0000
4.750 1.3579 0.01834 0.00998 -0.1662 0.4527 1.0000
5.000 1.3824 0.01860 0.01019 -0.1657 0.4505 1.0000
5.250 1.4078 0.01887 0.01040 -0.1654 0.4485 1.0000
5.500 1.4344 0.01918 0.01065 -0.1654 0.4467 1.0000
5.750 1.4556 0.01950 0.01103 -0.1643 0.4446 1.0000
6.000 1.4766 0.01982 0.01141 -0.1633 0.4422 1.0000
6.250 1.4977 0.02014 0.01176 -0.1623 0.4397 1.0000
6.500 1.5193 0.02045 0.01210 -0.1614 0.4373 1.0000
6.750 1.5413 0.02076 0.01242 -0.1606 0.4351 1.0000
7.000 1.5639 0.02106 0.01273 -0.1598 0.4332 1.0000
7.250 1.5881 0.02137 0.01301 -0.1594 0.4312 1.0000
7.500 1.6111 0.02171 0.01335 -0.1588 0.4291 1.0000
7.750 1.6256 0.02213 0.01388 -0.1568 0.4265 1.0000
8.000 1.6407 0.02253 0.01438 -0.1548 0.4239 1.0000
8.250 1.6567 0.02292 0.01482 -0.1530 0.4212 1.0000
8.500 1.6738 0.02328 0.01522 -0.1514 0.4186 1.0000
8.750 1.6925 0.02361 0.01557 -0.1501 0.4162 1.0000
9.000 1.7152 0.02391 0.01585 -0.1496 0.4139 1.0000
9.250 1.7265 0.02444 0.01648 -0.1472 0.4110 1.0000
9.500 1.7343 0.02505 0.01722 -0.1443 0.4077 1.0000
9.750 1.7454 0.02560 0.01786 -0.1420 0.4044 1.0000
10.000 1.7588 0.02608 0.01839 -0.1402 0.4014 1.0000
10.250 1.7763 0.02647 0.01878 -0.1389 0.3987 1.0000
10.500 1.7896 0.02706 0.01942 -0.1372 0.3958 1.0000
10.750 1.7907 0.02808 0.02062 -0.1339 0.3919 1.0000
11.000 1.7966 0.02898 0.02163 -0.1314 0.3883 1.0000
11.250 1.8064 0.02974 0.02244 -0.1295 0.3849 1.0000
11.500 1.8232 0.03022 0.02291 -0.1284 0.3819 1.0000
11.750 1.8203 0.03175 0.02462 -0.1254 0.3778 1.0000
12.000 1.8190 0.03333 0.02634 -0.1228 0.3734 1.0000
12.250 1.8245 0.03455 0.02762 -0.1210 0.3694 1.0000
12.500 1.8375 0.03532 0.02839 -0.1199 0.3657 1.0000
12.750 1.8241 0.03810 0.03138 -0.1171 0.3605 1.0000
13.000 1.8213 0.04022 0.03360 -0.1152 0.3554 1.0000
13.250 1.8325 0.04123 0.03460 -0.1142 0.3512 1.0000
13.500 1.8123 0.04521 0.03880 -0.1120 0.3450 1.0000
13.750 1.8095 0.04771 0.04137 -0.1108 0.3393 1.0000
14.000 1.8065 0.05036 0.04408 -0.1097 0.3339 1.0000
14.250 1.7875 0.05490 0.04877 -0.1085 0.3268 1.0000
14.500 1.7966 0.05634 0.05020 -0.1079 0.3215 1.0000
14.750 1.7654 0.06273 0.05678 -0.1071 0.3131 1.0000
15.000 1.7740 0.06431 0.05834 -0.1066 0.3071 1.0000
15.250 1.7440 0.07097 0.06515 -0.1063 0.2979 1.0000
15.500 1.7463 0.07347 0.06767 -0.1060 0.2913 1.0000
15.750 1.7272 0.07900 0.07330 -0.1061 0.2821 1.0000
16.000 1.7208 0.08283 0.07717 -0.1061 0.2743 1.0000
16.250 1.7166 0.08638 0.08075 -0.1061 0.2660 1.0000
16.500 1.7056 0.09102 0.08545 -0.1065 0.2573 1.0000
16.750 1.7105 0.09328 0.08768 -0.1064 0.2497 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 421 AIRFOIL (e421-il)