EPPLER 421 AIRFOIL (e421-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 421 AIRFOIL (e421-il) Reynolds number: 1,000,000 Max Cl/Cd: 141.05 at α=7.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e421-il-1000000.txt Download as CSV file: xf-e421-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 0.2181 0.08751 0.08384 -0.1339 0.6095 0.0162
-9.500 0.2260 0.08516 0.08148 -0.1346 0.6029 0.0162
-9.250 0.2292 0.08200 0.07830 -0.1355 0.5965 0.0162
-8.750 -0.0110 0.03744 0.03361 -0.1587 0.6076 0.0099
-8.500 -0.0069 0.02415 0.01981 -0.1775 0.6025 0.0097
-8.250 0.0163 0.01999 0.01530 -0.1819 0.5949 0.0098
-8.000 0.0407 0.01838 0.01347 -0.1830 0.5882 0.0098
-7.750 0.0657 0.01716 0.01206 -0.1836 0.5814 0.0099
-7.500 0.0906 0.01623 0.01095 -0.1838 0.5743 0.0101
-7.250 0.1167 0.01543 0.01000 -0.1840 0.5686 0.0102
-7.000 0.1430 0.01478 0.00922 -0.1841 0.5626 0.0105
-6.750 0.1691 0.01424 0.00854 -0.1839 0.5563 0.0106
-6.500 0.1956 0.01357 0.00772 -0.1840 0.5510 0.0110
-6.250 0.2222 0.01290 0.00694 -0.1840 0.5457 0.0115
-6.000 0.2490 0.01255 0.00649 -0.1838 0.5401 0.0119
-5.750 0.2759 0.01227 0.00613 -0.1836 0.5348 0.0124
-5.500 0.3037 0.01194 0.00573 -0.1835 0.5310 0.0128
-5.250 0.3313 0.01169 0.00539 -0.1834 0.5267 0.0132
-5.000 0.3584 0.01133 0.00494 -0.1832 0.5222 0.0139
-4.750 0.3854 0.01111 0.00464 -0.1830 0.5172 0.0147
-4.500 0.4137 0.01089 0.00437 -0.1829 0.5144 0.0157
-4.250 0.4416 0.01064 0.00407 -0.1828 0.5109 0.0171
-4.000 0.4692 0.01047 0.00385 -0.1826 0.5069 0.0194
-3.750 0.4965 0.01028 0.00362 -0.1824 0.5030 0.0234
-3.500 0.5236 0.01015 0.00346 -0.1821 0.4990 0.0296
-3.250 0.5518 0.00999 0.00332 -0.1821 0.4968 0.0376
-3.000 0.5798 0.00987 0.00322 -0.1820 0.4941 0.0458
-2.750 0.6076 0.00978 0.00314 -0.1818 0.4912 0.0539
-2.500 0.6350 0.00973 0.00308 -0.1816 0.4881 0.0625
-2.250 0.6621 0.00972 0.00305 -0.1814 0.4846 0.0715
-2.000 0.6892 0.00972 0.00304 -0.1811 0.4811 0.0807
-1.750 0.7172 0.00965 0.00301 -0.1810 0.4793 0.0908
-1.500 0.7451 0.00961 0.00300 -0.1809 0.4772 0.1016
-1.250 0.7728 0.00958 0.00300 -0.1807 0.4749 0.1122
-1.000 0.8002 0.00959 0.00300 -0.1805 0.4725 0.1213
-0.750 0.8273 0.00962 0.00302 -0.1803 0.4698 0.1299
-0.500 0.8541 0.00966 0.00306 -0.1799 0.4669 0.1398
-0.250 0.8808 0.00977 0.00312 -0.1796 0.4637 0.1464
0.000 0.9084 0.00975 0.00314 -0.1795 0.4623 0.1569
0.250 0.9359 0.00978 0.00317 -0.1793 0.4606 0.1634
0.500 0.9633 0.00978 0.00321 -0.1791 0.4587 0.1739
0.750 0.9904 0.00983 0.00326 -0.1789 0.4566 0.1807
1.000 1.0173 0.00987 0.00331 -0.1786 0.4545 0.1905
1.250 1.0437 0.00994 0.00337 -0.1783 0.4521 0.1980
1.500 1.0698 0.01003 0.00345 -0.1779 0.4495 0.2065
1.750 1.0958 0.01018 0.00357 -0.1776 0.4466 0.2152
2.000 1.1228 0.01020 0.00363 -0.1773 0.4455 0.2238
2.250 1.1496 0.01023 0.00370 -0.1771 0.4441 0.2337
2.500 1.1762 0.01028 0.00377 -0.1768 0.4424 0.2440
2.750 1.2026 0.01032 0.00385 -0.1766 0.4404 0.2551
3.000 1.2287 0.01038 0.00394 -0.1762 0.4385 0.2694
3.250 1.2544 0.01045 0.00404 -0.1758 0.4365 0.2851
3.500 1.2797 0.01054 0.00416 -0.1754 0.4345 0.3059
3.750 1.3047 0.01066 0.00431 -0.1749 0.4321 0.3325
4.000 1.3300 0.01081 0.00451 -0.1745 0.4296 0.3676
4.250 1.3555 0.01077 0.00464 -0.1742 0.4284 0.4319
4.500 1.3790 0.01005 0.00493 -0.1736 0.4270 1.0000
4.750 1.4037 0.01015 0.00502 -0.1730 0.4252 1.0000
5.000 1.4281 0.01027 0.00513 -0.1724 0.4235 1.0000
5.250 1.4521 0.01039 0.00525 -0.1717 0.4217 1.0000
5.500 1.4755 0.01053 0.00538 -0.1709 0.4197 1.0000
5.750 1.4983 0.01070 0.00552 -0.1700 0.4175 1.0000
6.000 1.5210 0.01092 0.00571 -0.1692 0.4149 1.0000
6.250 1.5445 0.01112 0.00591 -0.1685 0.4127 1.0000
6.500 1.5679 0.01122 0.00604 -0.1677 0.4112 1.0000
6.750 1.5910 0.01133 0.00618 -0.1670 0.4093 1.0000
7.000 1.6138 0.01146 0.00632 -0.1662 0.4070 1.0000
7.250 1.6362 0.01160 0.00648 -0.1653 0.4046 1.0000
7.500 1.6577 0.01179 0.00665 -0.1643 0.4018 1.0000
7.750 1.6780 0.01205 0.00689 -0.1631 0.3989 1.0000
8.000 1.6995 0.01227 0.00712 -0.1622 0.3963 1.0000
8.250 1.7218 0.01240 0.00730 -0.1614 0.3944 1.0000
8.500 1.7439 0.01256 0.00749 -0.1605 0.3920 1.0000
8.750 1.7654 0.01274 0.00769 -0.1596 0.3894 1.0000
9.000 1.7854 0.01297 0.00794 -0.1585 0.3868 1.0000
9.250 1.8034 0.01328 0.00823 -0.1570 0.3835 1.0000
9.500 1.8229 0.01356 0.00853 -0.1558 0.3805 1.0000
9.750 1.8441 0.01374 0.00877 -0.1550 0.3778 1.0000
10.000 1.8641 0.01397 0.00904 -0.1539 0.3747 1.0000
10.250 1.8823 0.01428 0.00936 -0.1526 0.3712 1.0000
10.500 1.8980 0.01469 0.00977 -0.1509 0.3676 1.0000
10.750 1.9159 0.01503 0.01015 -0.1497 0.3645 1.0000
11.000 1.9355 0.01531 0.01049 -0.1487 0.3612 1.0000
11.250 1.9521 0.01570 0.01091 -0.1473 0.3571 1.0000
11.500 1.9656 0.01624 0.01144 -0.1454 0.3530 1.0000
11.750 1.9810 0.01672 0.01196 -0.1439 0.3492 1.0000
12.000 1.9979 0.01715 0.01245 -0.1427 0.3452 1.0000
12.250 2.0100 0.01780 0.01311 -0.1409 0.3399 1.0000
12.500 2.0210 0.01855 0.01388 -0.1390 0.3349 1.0000
12.750 2.0351 0.01917 0.01455 -0.1375 0.3298 1.0000
13.000 2.0427 0.02015 0.01554 -0.1354 0.3236 1.0000
13.250 2.0526 0.02107 0.01649 -0.1337 0.3173 1.0000
13.500 2.0594 0.02221 0.01765 -0.1317 0.3106 1.0000
13.750 2.0642 0.02354 0.01899 -0.1296 0.3032 1.0000
14.000 2.0654 0.02519 0.02065 -0.1274 0.2947 1.0000
14.250 2.0686 0.02679 0.02228 -0.1255 0.2871 1.0000
14.500 2.0634 0.02911 0.02460 -0.1231 0.2784 1.0000
14.750 2.0625 0.03123 0.02675 -0.1213 0.2693 1.0000
15.000 2.0545 0.03405 0.02958 -0.1192 0.2603 1.0000
15.500 2.0289 0.04101 0.03656 -0.1154 0.2398 1.0000
15.750 2.0145 0.04491 0.04049 -0.1138 0.2311 1.0000
16.000 1.9971 0.04926 0.04486 -0.1124 0.2228 1.0000
16.250 1.9845 0.05326 0.04890 -0.1114 0.2149 1.0000
16.500 1.9627 0.05843 0.05409 -0.1104 0.2067 1.0000
16.750 1.9496 0.06275 0.05846 -0.1098 0.1993 1.0000
17.000 1.9305 0.06793 0.06367 -0.1093 0.1919 1.0000
17.250 1.9174 0.07246 0.06823 -0.1091 0.1849 1.0000
17.500 1.9014 0.07747 0.07327 -0.1090 0.1782 1.0000
17.750 1.8881 0.08225 0.07810 -0.1091 0.1715 1.0000
18.000 1.8738 0.08725 0.08312 -0.1093 0.1652 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 421 AIRFOIL (e421-il)