EPPLER 435 AIRFOIL (e435-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 435 AIRFOIL (e435-il) Reynolds number: 1,000,000 Max Cl/Cd: 132.96 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e435-il-1000000-n5.txt Download as CSV file: xf-e435-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 435 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.250 -0.3778 0.05814 0.05578 -0.1304 0.9375 0.0038
-13.000 -0.3820 0.04455 0.04178 -0.1502 0.9264 0.0038
-12.750 -0.3714 0.03770 0.03453 -0.1622 0.9020 0.0038
-12.500 -0.3760 0.03412 0.03064 -0.1649 0.8733 0.0038
-12.250 -0.3864 0.03129 0.02754 -0.1647 0.8527 0.0038
-12.000 -0.3927 0.02906 0.02510 -0.1639 0.8378 0.0038
-11.750 -0.3950 0.02722 0.02308 -0.1629 0.8260 0.0038
-11.500 -0.3947 0.02567 0.02137 -0.1616 0.8154 0.0038
-11.250 -0.3937 0.02416 0.01971 -0.1601 0.8064 0.0038
-11.000 -0.3905 0.02290 0.01830 -0.1585 0.7990 0.0038
-10.750 -0.3868 0.02172 0.01699 -0.1567 0.7920 0.0038
-10.250 -0.3807 0.01963 0.01462 -0.1521 0.7792 0.0038
-10.000 -0.3685 0.01871 0.01357 -0.1508 0.7737 0.0038
-9.500 -0.3364 0.01717 0.01180 -0.1489 0.7651 0.0038
-9.250 -0.3179 0.01644 0.01097 -0.1481 0.7609 0.0038
-9.000 -0.2982 0.01582 0.01025 -0.1474 0.7565 0.0038
-8.500 -0.2559 0.01464 0.00890 -0.1464 0.7494 0.0038
-8.250 -0.2333 0.01409 0.00828 -0.1461 0.7460 0.0038
-8.000 -0.2100 0.01362 0.00774 -0.1459 0.7424 0.0039
-7.750 -0.1872 0.01305 0.00709 -0.1455 0.7388 0.0039
-7.500 -0.1632 0.01260 0.00657 -0.1453 0.7355 0.0039
-7.250 -0.1382 0.01213 0.00604 -0.1453 0.7328 0.0040
-7.000 -0.1128 0.01169 0.00556 -0.1454 0.7297 0.0041
-6.750 -0.0867 0.01135 0.00517 -0.1454 0.7264 0.0042
-6.500 -0.0604 0.01105 0.00483 -0.1455 0.7232 0.0044
-6.250 -0.0338 0.01079 0.00452 -0.1456 0.7202 0.0046
-6.000 -0.0064 0.01054 0.00424 -0.1459 0.7176 0.0047
-5.750 0.0213 0.01029 0.00397 -0.1461 0.7148 0.0051
-5.500 0.0491 0.01008 0.00372 -0.1464 0.7117 0.0055
-5.250 0.0768 0.00987 0.00349 -0.1467 0.7086 0.0062
-5.000 0.1046 0.00968 0.00327 -0.1469 0.7056 0.0076
-4.750 0.1326 0.00950 0.00308 -0.1472 0.7027 0.0101
-4.500 0.1611 0.00930 0.00290 -0.1477 0.7001 0.0148
-4.250 0.1897 0.00911 0.00274 -0.1481 0.6971 0.0223
-4.000 0.2182 0.00895 0.00260 -0.1485 0.6940 0.0313
-3.750 0.2466 0.00878 0.00246 -0.1489 0.6908 0.0442
-3.500 0.2751 0.00854 0.00232 -0.1494 0.6879 0.0748
-3.250 0.3043 0.00821 0.00218 -0.1502 0.6851 0.1266
-3.000 0.3341 0.00778 0.00202 -0.1512 0.6821 0.2088
-2.750 0.3638 0.00746 0.00189 -0.1521 0.6788 0.2775
-2.500 0.3939 0.00705 0.00175 -0.1532 0.6755 0.3728
-2.250 0.4241 0.00663 0.00166 -0.1543 0.6722 0.4958
-2.000 0.4534 0.00657 0.00167 -0.1548 0.6692 0.5301
-1.750 0.4826 0.00656 0.00167 -0.1552 0.6658 0.5474
-1.500 0.5115 0.00658 0.00167 -0.1555 0.6620 0.5588
-1.250 0.5400 0.00660 0.00169 -0.1558 0.6581 0.5701
-1.000 0.5683 0.00664 0.00172 -0.1560 0.6544 0.5822
-0.750 0.5973 0.00667 0.00174 -0.1563 0.6505 0.5933
-0.500 0.6257 0.00670 0.00178 -0.1565 0.6461 0.6015
-0.250 0.6537 0.00676 0.00180 -0.1567 0.6416 0.6053
0.000 0.6819 0.00681 0.00181 -0.1568 0.6375 0.6075
0.250 0.7103 0.00684 0.00183 -0.1571 0.6327 0.6097
0.500 0.7380 0.00688 0.00186 -0.1572 0.6273 0.6118
0.750 0.7654 0.00695 0.00190 -0.1572 0.6219 0.6139
1.000 0.7934 0.00699 0.00194 -0.1574 0.6161 0.6163
1.250 0.8205 0.00706 0.00199 -0.1574 0.6100 0.6188
1.500 0.8479 0.00713 0.00204 -0.1575 0.6044 0.6210
1.750 0.8753 0.00720 0.00209 -0.1575 0.5979 0.6231
2.000 0.9015 0.00730 0.00215 -0.1573 0.5911 0.6249
2.250 0.9287 0.00736 0.00222 -0.1574 0.5838 0.6270
2.500 0.9542 0.00747 0.00231 -0.1571 0.5756 0.6292
2.750 0.9807 0.00756 0.00239 -0.1570 0.5670 0.6316
3.000 1.0059 0.00769 0.00249 -0.1566 0.5583 0.6342
3.250 1.0313 0.00781 0.00260 -0.1563 0.5482 0.6368
3.500 1.0561 0.00795 0.00271 -0.1559 0.5381 0.6391
3.750 1.0796 0.00812 0.00284 -0.1552 0.5260 0.6412
4.000 1.1024 0.00831 0.00299 -0.1544 0.5127 0.6434
4.250 1.1253 0.00850 0.00315 -0.1536 0.4989 0.6457
4.500 1.1473 0.00870 0.00332 -0.1527 0.4849 0.6482
4.750 1.1675 0.00893 0.00350 -0.1513 0.4702 0.6507
5.000 1.1861 0.00917 0.00370 -0.1497 0.4547 0.6535
5.250 1.2036 0.00946 0.00393 -0.1479 0.4380 0.6563
5.500 1.2206 0.00979 0.00419 -0.1460 0.4205 0.6587
5.750 1.2363 0.01018 0.00451 -0.1439 0.4011 0.6614
6.000 1.2519 0.01057 0.00483 -0.1418 0.3816 0.6642
6.250 1.2675 0.01097 0.00518 -0.1398 0.3631 0.6670
6.500 1.2827 0.01140 0.00555 -0.1377 0.3453 0.6699
6.750 1.2967 0.01187 0.00595 -0.1355 0.3268 0.6727
7.000 1.3106 0.01237 0.00638 -0.1333 0.3096 0.6753
7.250 1.3241 0.01288 0.00684 -0.1311 0.2931 0.6782
7.500 1.3377 0.01340 0.00733 -0.1289 0.2777 0.6814
7.750 1.3504 0.01399 0.00787 -0.1267 0.2618 0.6850
8.000 1.3626 0.01461 0.00844 -0.1245 0.2466 0.6886
8.250 1.3734 0.01533 0.00909 -0.1221 0.2301 0.6918
8.500 1.3851 0.01602 0.00976 -0.1199 0.2159 0.6951
8.750 1.3951 0.01683 0.01052 -0.1176 0.2004 0.6987
9.000 1.4055 0.01766 0.01131 -0.1154 0.1862 0.7024
9.250 1.4149 0.01858 0.01218 -0.1132 0.1714 0.7062
9.750 1.4359 0.02041 0.01398 -0.1093 0.1479 0.7139
10.000 1.4449 0.02147 0.01500 -0.1073 0.1353 0.7187
10.250 1.4539 0.02257 0.01607 -0.1054 0.1238 0.7236
10.500 1.4625 0.02373 0.01721 -0.1035 0.1131 0.7283
10.750 1.4719 0.02488 0.01836 -0.1018 0.1033 0.7336
11.000 1.4820 0.02601 0.01950 -0.1003 0.0951 0.7390
11.250 1.4907 0.02728 0.02076 -0.0987 0.0873 0.7443
11.500 1.4985 0.02865 0.02212 -0.0971 0.0785 0.7504
11.750 1.5065 0.03004 0.02352 -0.0956 0.0699 0.7570
12.000 1.5146 0.03147 0.02496 -0.0942 0.0630 0.7643
12.250 1.5209 0.03308 0.02657 -0.0927 0.0550 0.7732
12.500 1.5288 0.03460 0.02812 -0.0914 0.0495 0.7826
12.750 1.5343 0.03636 0.02989 -0.0900 0.0423 0.7934
13.000 1.5417 0.03799 0.03156 -0.0889 0.0373 0.8064
13.250 1.5473 0.03983 0.03344 -0.0877 0.0320 0.8221
13.500 1.5536 0.04160 0.03529 -0.0866 0.0282 0.8446
13.750 1.5577 0.04313 0.03701 -0.0848 0.0254 0.9161
14.000 1.5638 0.04492 0.03888 -0.0839 0.0229 1.0000
14.250 1.5698 0.04694 0.04091 -0.0831 0.0202 1.0000
14.500 1.5770 0.04888 0.04290 -0.0824 0.0185 1.0000
14.750 1.5824 0.05106 0.04510 -0.0817 0.0162 1.0000
15.000 1.5880 0.05327 0.04735 -0.0812 0.0144 1.0000
15.250 1.5932 0.05554 0.04966 -0.0807 0.0128 1.0000
15.500 1.5974 0.05798 0.05214 -0.0803 0.0112 1.0000
15.750 1.6025 0.06037 0.05459 -0.0799 0.0104 1.0000
16.000 1.6069 0.06288 0.05714 -0.0797 0.0092 1.0000
16.250 1.6104 0.06556 0.05989 -0.0795 0.0085 1.0000
16.500 1.6147 0.06817 0.06256 -0.0794 0.0078 1.0000
16.750 1.6172 0.07106 0.06552 -0.0794 0.0070 1.0000
17.000 1.6194 0.07405 0.06857 -0.0795 0.0064 1.0000
17.250 1.6221 0.07702 0.07161 -0.0797 0.0059 1.0000
17.500 1.6228 0.08032 0.07498 -0.0800 0.0053 1.0000
17.750 1.6234 0.08367 0.07840 -0.0804 0.0049 1.0000
18.000 1.6241 0.08704 0.08184 -0.0809 0.0045 1.0000
18.250 1.6238 0.09062 0.08551 -0.0815 0.0041 1.0000
18.500 1.6221 0.09447 0.08944 -0.0823 0.0038 1.0000
18.750 1.6210 0.09827 0.09333 -0.0832 0.0035 1.0000
19.000 1.6200 0.10208 0.09722 -0.0842 0.0033 1.0000
19.250 1.6175 0.10615 0.10138 -0.0854 0.0030 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 435 AIRFOIL (e435-il)