EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: EPPLER 59 AIRFOIL (e59-il) Reynolds number: 500,000 Max Cl/Cd: 119.29 at α=0.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-e59-il-500000-n5.txt Download as CSV file: xf-e59-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 59 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.1660 0.08996 0.08763 -0.0684 0.9766 0.0085
-9.250 -0.1617 0.08655 0.08422 -0.0690 0.9746 0.0084
-9.000 -0.1544 0.08305 0.08072 -0.0694 0.9732 0.0077
-8.750 -0.1478 0.07934 0.07701 -0.0709 0.9720 0.0074
-8.500 -0.2355 0.09511 0.09267 -0.0655 0.9793 0.0079
-8.250 -0.2275 0.09213 0.08968 -0.0670 0.9775 0.0071
-8.000 -0.2222 0.08911 0.08667 -0.0680 0.9740 0.0066
-7.750 -0.2135 0.08491 0.08249 -0.0711 0.9715 0.0061
-7.500 -0.1995 0.08154 0.07912 -0.0746 0.9698 0.0063
-7.250 -0.1945 0.07870 0.07630 -0.0757 0.9656 0.0064
-7.000 -0.1856 0.07531 0.07292 -0.0782 0.9609 0.0065
-6.750 -0.1640 0.07107 0.06868 -0.0842 0.9586 0.0073
-5.750 0.0260 0.01817 0.01423 -0.1656 0.9457 0.0074
-5.500 0.0673 0.01501 0.01040 -0.1703 0.9455 0.0081
-5.250 0.1040 0.01344 0.00845 -0.1729 0.9452 0.0088
-5.000 0.1382 0.01276 0.00765 -0.1746 0.9446 0.0102
-4.750 0.1727 0.01207 0.00679 -0.1763 0.9442 0.0113
-4.500 0.1958 0.01156 0.00614 -0.1754 0.9403 0.0125
-4.250 0.2258 0.01090 0.00532 -0.1760 0.9381 0.0145
-4.000 0.2557 0.01067 0.00509 -0.1765 0.9363 0.0184
-3.750 0.2872 0.01040 0.00474 -0.1773 0.9349 0.0216
-3.500 0.3197 0.01007 0.00438 -0.1784 0.9338 0.0259
-3.250 0.3521 0.00988 0.00415 -0.1794 0.9326 0.0300
-3.000 0.3848 0.00970 0.00391 -0.1804 0.9317 0.0324
-2.750 0.4186 0.00935 0.00355 -0.1819 0.9309 0.0372
-2.500 0.4524 0.00913 0.00331 -0.1832 0.9302 0.0407
-2.250 0.4733 0.00905 0.00319 -0.1816 0.9252 0.0426
-2.000 0.5021 0.00888 0.00299 -0.1818 0.9221 0.0443
-1.750 0.5346 0.00863 0.00274 -0.1828 0.9200 0.0492
-1.500 0.5680 0.00844 0.00256 -0.1840 0.9182 0.0557
-1.250 0.6045 0.00818 0.00232 -0.1858 0.9160 0.0705
-1.000 0.6315 0.00790 0.00210 -0.1855 0.9047 0.0936
-0.750 0.6673 0.00751 0.00185 -0.1871 0.8908 0.1535
-0.500 0.7037 0.00726 0.00163 -0.1889 0.8757 0.1914
-0.250 0.7386 0.00704 0.00152 -0.1904 0.8609 0.2483
0.000 0.7742 0.00676 0.00144 -0.1922 0.8307 0.3857
0.250 0.8040 0.00674 0.00147 -0.1926 0.7723 0.5374
0.500 0.8245 0.00704 0.00164 -0.1909 0.7014 0.6362
0.750 0.8356 0.00769 0.00203 -0.1872 0.5794 0.7751
1.000 0.8441 0.00791 0.00224 -0.1827 0.5266 0.9270
1.250 0.8671 0.00822 0.00239 -0.1817 0.4939 1.0000
1.500 0.8909 0.00859 0.00256 -0.1810 0.4554 1.0000
1.750 0.9076 0.00964 0.00294 -0.1790 0.3164 1.0000
2.000 0.9290 0.01036 0.00330 -0.1779 0.2445 1.0000
2.250 0.9522 0.01090 0.00356 -0.1771 0.1905 1.0000
2.500 0.9714 0.01195 0.00405 -0.1757 0.0807 1.0000
2.750 0.9953 0.01243 0.00438 -0.1750 0.0458 1.0000
3.000 1.0205 0.01273 0.00463 -0.1745 0.0405 1.0000
3.250 1.0460 0.01298 0.00489 -0.1741 0.0384 1.0000
3.500 1.0713 0.01324 0.00516 -0.1736 0.0376 1.0000
3.750 1.0964 0.01351 0.00547 -0.1730 0.0368 1.0000
4.000 1.1213 0.01381 0.00580 -0.1725 0.0363 1.0000
4.250 1.1460 0.01412 0.00615 -0.1719 0.0359 1.0000
4.500 1.1705 0.01445 0.00654 -0.1712 0.0355 1.0000
4.750 1.1946 0.01482 0.00697 -0.1706 0.0351 1.0000
5.000 1.2184 0.01521 0.00742 -0.1698 0.0343 1.0000
5.250 1.2416 0.01567 0.00791 -0.1690 0.0322 1.0000
5.500 1.2637 0.01626 0.00855 -0.1680 0.0295 1.0000
5.750 1.2859 0.01682 0.00919 -0.1670 0.0277 1.0000
6.000 1.3105 0.01702 0.00943 -0.1664 0.0266 1.0000
6.250 1.3345 0.01728 0.00975 -0.1658 0.0241 1.0000
6.500 1.3582 0.01756 0.01006 -0.1651 0.0197 1.0000
6.750 1.3809 0.01798 0.01048 -0.1643 0.0159 1.0000
7.000 1.4029 0.01848 0.01091 -0.1633 0.0117 1.0000
7.250 1.4235 0.01913 0.01162 -0.1620 0.0094 1.0000
7.500 1.4430 0.01990 0.01244 -0.1606 0.0077 1.0000
7.750 1.4620 0.02074 0.01338 -0.1590 0.0069 1.0000
8.000 1.4810 0.02155 0.01431 -0.1574 0.0063 1.0000
8.250 1.4989 0.02242 0.01527 -0.1557 0.0057 1.0000
8.500 1.5156 0.02335 0.01630 -0.1538 0.0053 1.0000
8.750 1.5289 0.02469 0.01775 -0.1514 0.0048 1.0000
9.000 1.5448 0.02558 0.01878 -0.1494 0.0045 1.0000
9.250 1.5594 0.02662 0.01996 -0.1472 0.0042 1.0000
9.500 1.5733 0.02772 0.02118 -0.1449 0.0039 1.0000
9.750 1.5860 0.02893 0.02253 -0.1425 0.0037 1.0000
10.000 1.5979 0.03019 0.02392 -0.1401 0.0035 1.0000
10.250 1.6082 0.03163 0.02555 -0.1375 0.0034 1.0000
10.500 1.6160 0.03333 0.02741 -0.1346 0.0032 1.0000
10.750 1.6191 0.03563 0.02992 -0.1312 0.0031 1.0000
11.000 1.6256 0.03750 0.03201 -0.1283 0.0030 1.0000
11.250 1.6298 0.03962 0.03436 -0.1253 0.0030 1.0000
11.500 1.6326 0.04189 0.03688 -0.1223 0.0029 1.0000
11.750 1.6314 0.04464 0.03989 -0.1190 0.0028 1.0000
12.000 1.6281 0.04761 0.04312 -0.1159 0.0027 1.0000
12.250 1.6221 0.05096 0.04675 -0.1129 0.0027 1.0000
12.500 1.6126 0.05477 0.05084 -0.1100 0.0026 1.0000
12.750 1.6012 0.05892 0.05525 -0.1077 0.0026 1.0000
13.000 1.5871 0.06365 0.06025 -0.1059 0.0026 1.0000
13.250 1.5704 0.06900 0.06586 -0.1048 0.0025 1.0000
13.500 1.5506 0.07524 0.07237 -0.1049 0.0025 1.0000
13.750 1.5302 0.08211 0.07947 -0.1061 0.0026 1.0000
14.000 1.5072 0.09033 0.08793 -0.1091 0.0026 1.0000
14.250 1.4849 0.09949 0.09731 -0.1137 0.0025 1.0000
14.500 1.4615 0.11006 0.10808 -0.1200 0.0026 1.0000
14.750 1.4363 0.12242 0.12064 -0.1280 0.0026 1.0000
15.000 1.4121 0.13616 0.13454 -0.1370 0.0027 1.0000
15.250 1.3819 0.15258 0.15110 -0.1474 0.0028 1.0000
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Polar data table (+)
Polar graphs
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