EPPLER 479 AIRFOIL (e479-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: EPPLER 479 AIRFOIL (e479-il) Reynolds number: 100,000 Max Cl/Cd: 33.08 at α=9.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-e479-il-100000.txt Download as CSV file: xf-e479-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: EPPLER 479 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-14.250 -1.0729 0.06284 0.05659 -0.0362 1.0000 0.1489
-14.000 -1.1318 0.05491 0.04836 -0.0386 1.0000 0.1491
-13.750 -1.1736 0.05007 0.04319 -0.0379 1.0000 0.1501
-13.500 -1.1189 0.05145 0.04490 -0.0375 1.0000 0.1578
-13.250 -1.1435 0.04796 0.04120 -0.0363 1.0000 0.1606
-13.000 -1.1713 0.04494 0.03791 -0.0334 1.0000 0.1630
-12.750 -1.1899 0.04263 0.03535 -0.0296 1.0000 0.1660
-12.500 -1.1522 0.04263 0.03562 -0.0302 1.0000 0.1731
-12.250 -1.1645 0.04064 0.03337 -0.0268 1.0000 0.1775
-12.000 -1.1539 0.03923 0.03193 -0.0253 1.0000 0.1829
-11.750 -1.1368 0.03858 0.03132 -0.0241 1.0000 0.1892
-11.500 -1.1453 0.03665 0.02897 -0.0207 1.0000 0.1946
-11.250 -1.1131 0.03635 0.02896 -0.0209 1.0000 0.2014
-11.000 -1.1089 0.03501 0.02738 -0.0185 1.0000 0.2081
-10.750 -1.0870 0.03429 0.02676 -0.0177 1.0000 0.2147
-10.500 -1.0735 0.03345 0.02581 -0.0160 1.0000 0.2220
-10.250 -1.0566 0.03256 0.02488 -0.0147 1.0000 0.2291
-10.000 -1.0378 0.03201 0.02431 -0.0135 1.0000 0.2367
-9.750 -1.0222 0.03108 0.02329 -0.0120 1.0000 0.2442
-9.500 -1.0010 0.03069 0.02293 -0.0109 1.0000 0.2520
-9.250 -0.9852 0.02978 0.02190 -0.0093 1.0000 0.2598
-9.000 -0.9629 0.02946 0.02164 -0.0083 1.0000 0.2678
-8.750 -0.9463 0.02863 0.02071 -0.0068 1.0000 0.2759
-8.500 -0.9240 0.02831 0.02043 -0.0057 1.0000 0.2839
-8.250 -0.9067 0.02756 0.01959 -0.0042 1.0000 0.2923
-8.000 -0.8847 0.02724 0.01932 -0.0031 1.0000 0.3004
-7.750 -0.8671 0.02653 0.01853 -0.0016 1.0000 0.3088
-7.500 -0.8454 0.02622 0.01827 -0.0004 1.0000 0.3170
-7.250 -0.8279 0.02555 0.01753 0.0012 1.0000 0.3255
-7.000 -0.8066 0.02524 0.01728 0.0024 1.0000 0.3336
-6.750 -0.7894 0.02463 0.01659 0.0041 1.0000 0.3422
-6.500 -0.7690 0.02433 0.01637 0.0054 1.0000 0.3502
-6.250 -0.7535 0.02374 0.01571 0.0072 1.0000 0.3588
-6.000 -0.7350 0.02350 0.01557 0.0088 1.0000 0.3664
-5.750 -0.7235 0.02297 0.01493 0.0111 1.0000 0.3750
-5.500 -0.7087 0.02279 0.01489 0.0131 1.0000 0.3817
-5.250 -0.7023 0.02241 0.01433 0.0157 1.0000 0.3905
-5.000 -0.6797 0.02232 0.01444 0.0159 0.9971 0.3975
-4.750 -0.6229 0.02205 0.01411 0.0099 0.9818 0.4107
-4.500 -0.5679 0.02179 0.01381 0.0044 0.9663 0.4237
-4.250 -0.5131 0.02151 0.01365 -0.0005 0.9509 0.4345
-4.000 -0.4601 0.02107 0.01318 -0.0051 0.9351 0.4466
-3.750 -0.4147 0.02067 0.01270 -0.0083 0.9158 0.4591
-3.500 -0.3711 0.02036 0.01250 -0.0105 0.8948 0.4685
-3.250 -0.3352 0.02000 0.01203 -0.0116 0.8719 0.4795
-3.000 -0.3030 0.01981 0.01183 -0.0115 0.8489 0.4892
-2.750 -0.2770 0.01960 0.01155 -0.0106 0.8235 0.4987
-2.500 -0.2512 0.01946 0.01133 -0.0096 0.8006 0.5087
-2.250 -0.2256 0.01932 0.01113 -0.0085 0.7795 0.5180
-2.000 -0.2020 0.01920 0.01088 -0.0074 0.7575 0.5291
-1.750 -0.1765 0.01912 0.01083 -0.0063 0.7370 0.5381
-1.500 -0.1521 0.01901 0.01062 -0.0053 0.7178 0.5495
-1.250 -0.1268 0.01896 0.01056 -0.0043 0.6995 0.5600
-1.000 -0.1015 0.01889 0.01046 -0.0034 0.6824 0.5711
-0.750 -0.0766 0.01884 0.01031 -0.0025 0.6662 0.5840
-0.500 -0.0508 0.01882 0.01030 -0.0016 0.6510 0.5955
-0.250 -0.0251 0.01879 0.01023 -0.0008 0.6368 0.6080
0.000 0.0000 0.01877 0.01022 0.0000 0.6219 0.6219
0.250 0.0251 0.01879 0.01023 0.0008 0.6080 0.6368
0.500 0.0508 0.01882 0.01030 0.0016 0.5955 0.6510
0.750 0.0766 0.01884 0.01031 0.0025 0.5840 0.6662
1.000 0.1015 0.01889 0.01045 0.0034 0.5711 0.6824
1.250 0.1268 0.01896 0.01056 0.0043 0.5600 0.6995
1.500 0.1521 0.01901 0.01062 0.0053 0.5495 0.7178
1.750 0.1765 0.01912 0.01084 0.0063 0.5381 0.7370
2.000 0.2020 0.01920 0.01088 0.0074 0.5291 0.7575
2.250 0.2256 0.01932 0.01113 0.0085 0.5180 0.7795
2.500 0.2512 0.01946 0.01133 0.0096 0.5087 0.8006
2.750 0.2770 0.01960 0.01155 0.0106 0.4986 0.8235
3.000 0.3030 0.01981 0.01183 0.0115 0.4892 0.8489
3.250 0.3352 0.02000 0.01203 0.0116 0.4795 0.8719
3.500 0.3711 0.02036 0.01250 0.0105 0.4685 0.8947
3.750 0.4146 0.02067 0.01270 0.0083 0.4591 0.9158
4.000 0.4601 0.02107 0.01318 0.0051 0.4467 0.9352
4.250 0.5130 0.02152 0.01365 0.0005 0.4346 0.9510
4.500 0.5679 0.02179 0.01381 -0.0044 0.4237 0.9663
4.750 0.6230 0.02205 0.01410 -0.0099 0.4106 0.9819
5.000 0.6798 0.02231 0.01443 -0.0159 0.3974 0.9972
5.250 0.7021 0.02240 0.01433 -0.0157 0.3906 1.0000
5.500 0.7086 0.02279 0.01489 -0.0131 0.3817 1.0000
5.750 0.7233 0.02297 0.01494 -0.0111 0.3750 1.0000
6.000 0.7349 0.02350 0.01557 -0.0088 0.3665 1.0000
6.250 0.7534 0.02374 0.01572 -0.0072 0.3588 1.0000
6.500 0.7689 0.02433 0.01636 -0.0054 0.3502 1.0000
6.750 0.7893 0.02462 0.01659 -0.0040 0.3422 1.0000
7.000 0.8066 0.02524 0.01728 -0.0024 0.3336 1.0000
7.250 0.8278 0.02556 0.01753 -0.0012 0.3255 1.0000
7.500 0.8454 0.02622 0.01827 0.0004 0.3170 1.0000
7.750 0.8671 0.02654 0.01854 0.0016 0.3088 1.0000
8.000 0.8847 0.02724 0.01931 0.0031 0.3004 1.0000
8.250 0.9067 0.02757 0.01960 0.0042 0.2923 1.0000
8.500 0.9240 0.02831 0.02043 0.0057 0.2839 1.0000
8.750 0.9463 0.02863 0.02070 0.0068 0.2759 1.0000
9.000 0.9630 0.02946 0.02163 0.0083 0.2678 1.0000
9.250 0.9852 0.02978 0.02191 0.0093 0.2598 1.0000
9.500 1.0011 0.03069 0.02293 0.0109 0.2520 1.0000
9.750 1.0223 0.03108 0.02329 0.0119 0.2442 1.0000
10.000 1.0380 0.03201 0.02431 0.0134 0.2367 1.0000
10.250 1.0567 0.03255 0.02487 0.0147 0.2291 1.0000
10.500 1.0738 0.03345 0.02580 0.0160 0.2221 1.0000
10.750 1.0871 0.03430 0.02676 0.0177 0.2148 1.0000
11.000 1.1091 0.03501 0.02738 0.0185 0.2081 1.0000
11.250 1.1134 0.03634 0.02895 0.0209 0.2014 1.0000
11.500 1.1455 0.03665 0.02897 0.0207 0.1946 1.0000
11.750 1.1371 0.03858 0.03132 0.0241 0.1891 1.0000
12.000 1.1546 0.03922 0.03191 0.0252 0.1828 1.0000
12.250 1.1646 0.04063 0.03337 0.0268 0.1774 1.0000
12.500 1.1528 0.04263 0.03562 0.0301 0.1731 1.0000
12.750 1.1928 0.04259 0.03528 0.0294 0.1658 1.0000
13.000 1.1716 0.04496 0.03793 0.0333 0.1630 1.0000
13.250 1.1440 0.04796 0.04120 0.0363 0.1604 1.0000
13.500 1.1201 0.05143 0.04487 0.0374 0.1577 1.0000
13.750 1.1736 0.05010 0.04323 0.0378 0.1501 1.0000
14.000 1.1321 0.05496 0.04841 0.0384 0.1490 1.0000
14.250 1.0725 0.06300 0.05676 0.0359 0.1489 1.0000
|
Polar data table (+)
Polar graphs
<< Back to EPPLER 479 AIRFOIL (e479-il)