GOE 441 AIRFOIL (goe441-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 441 AIRFOIL (goe441-il) Reynolds number: 100,000 Max Cl/Cd: 51.95 at α=4° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe441-il-100000-n5.txt Download as CSV file: xf-goe441-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 441 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 0.0522 0.11190 0.10656 -0.0972 0.9081 0.0789
-9.750 0.0571 0.10934 0.10398 -0.0999 0.8989 0.0792
-9.500 0.0627 0.10664 0.10127 -0.1023 0.8898 0.0793
-9.250 0.0690 0.10386 0.09847 -0.1042 0.8804 0.0794
-9.000 0.0777 0.10086 0.09545 -0.1058 0.8712 0.0794
-8.750 0.0862 0.09799 0.09257 -0.1070 0.8618 0.0794
-7.750 0.1250 0.08405 0.07848 -0.1102 0.8269 0.0635
-7.500 0.1357 0.08157 0.07598 -0.1106 0.8172 0.0623
-7.250 0.1421 0.07880 0.07319 -0.1115 0.8074 0.0610
-6.750 0.1354 0.07030 0.06464 -0.1174 0.7875 0.0573
-6.500 0.1460 0.06717 0.06145 -0.1198 0.7782 0.0571
-6.250 0.1546 0.06368 0.05795 -0.1231 0.7673 0.0570
-6.000 0.1660 0.05680 0.05097 -0.1327 0.7584 0.0575
-5.500 0.2253 0.03744 0.03086 -0.1650 0.7383 0.0589
-5.250 0.2623 0.03214 0.02496 -0.1740 0.7253 0.0603
-5.000 0.2999 0.02873 0.02092 -0.1794 0.7131 0.0624
-4.750 0.3343 0.02677 0.01849 -0.1822 0.7004 0.0650
-4.500 0.3619 0.02615 0.01777 -0.1827 0.6852 0.0669
-4.250 0.3923 0.02498 0.01627 -0.1839 0.6697 0.0694
-4.000 0.4229 0.02387 0.01483 -0.1849 0.6537 0.0720
-3.750 0.4510 0.02346 0.01431 -0.1851 0.6372 0.0746
-3.500 0.4811 0.02280 0.01330 -0.1857 0.6217 0.0796
-3.250 0.5094 0.02254 0.01295 -0.1859 0.6078 0.0834
-3.000 0.5390 0.02205 0.01218 -0.1863 0.5953 0.0890
-2.750 0.5676 0.02183 0.01181 -0.1866 0.5846 0.0951
-2.500 0.5966 0.02155 0.01135 -0.1869 0.5752 0.1024
-2.250 0.6256 0.02133 0.01095 -0.1872 0.5665 0.1113
-2.000 0.6544 0.02118 0.01068 -0.1875 0.5590 0.1206
-1.750 0.6832 0.02106 0.01051 -0.1878 0.5525 0.1300
-1.500 0.7119 0.02100 0.01039 -0.1881 0.5463 0.1400
-1.250 0.7409 0.02103 0.01029 -0.1883 0.5408 0.1508
-1.000 0.7699 0.02109 0.01029 -0.1886 0.5357 0.1615
-0.750 0.7978 0.02116 0.01035 -0.1887 0.5301 0.1727
-0.500 0.8262 0.02124 0.01037 -0.1889 0.5250 0.1855
-0.250 0.8557 0.02132 0.01036 -0.1892 0.5206 0.2001
0.000 0.8854 0.02141 0.01040 -0.1897 0.5167 0.2182
0.250 0.9135 0.02146 0.01053 -0.1899 0.5122 0.2402
0.500 0.9420 0.02154 0.01068 -0.1902 0.5078 0.2722
0.750 0.9709 0.02165 0.01091 -0.1906 0.5036 0.3198
1.000 1.0003 0.02186 0.01116 -0.1909 0.4999 0.3829
1.250 1.0278 0.02213 0.01146 -0.1909 0.4962 0.4374
1.500 1.0539 0.02239 0.01181 -0.1906 0.4924 0.4815
1.750 1.0801 0.02266 0.01213 -0.1904 0.4884 0.5222
2.000 1.1066 0.02290 0.01240 -0.1901 0.4846 0.5616
2.250 1.1336 0.02314 0.01266 -0.1900 0.4812 0.6027
2.500 1.1609 0.02339 0.01293 -0.1898 0.4782 0.6499
2.750 1.1823 0.02359 0.01335 -0.1886 0.4747 0.7101
3.000 1.2010 0.02362 0.01364 -0.1867 0.4711 0.8102
3.250 1.2236 0.02383 0.01386 -0.1858 0.4674 1.0000
3.500 1.2507 0.02425 0.01417 -0.1859 0.4639 1.0000
3.750 1.2788 0.02467 0.01445 -0.1861 0.4608 1.0000
4.000 1.3061 0.02514 0.01481 -0.1862 0.4579 1.0000
4.250 1.3275 0.02566 0.01539 -0.1854 0.4541 1.0000
4.500 1.3503 0.02616 0.01591 -0.1847 0.4503 1.0000
4.750 1.3742 0.02663 0.01634 -0.1843 0.4467 1.0000
5.000 1.3996 0.02706 0.01671 -0.1840 0.4435 1.0000
5.250 1.4267 0.02748 0.01704 -0.1841 0.4407 1.0000
5.500 1.4483 0.02805 0.01763 -0.1833 0.4374 1.0000
5.750 1.4660 0.02868 0.01836 -0.1819 0.4334 1.0000
6.000 1.4861 0.02925 0.01898 -0.1809 0.4297 1.0000
6.250 1.5083 0.02974 0.01946 -0.1802 0.4264 1.0000
6.500 1.5330 0.03015 0.01983 -0.1799 0.4234 1.0000
6.750 1.5592 0.03058 0.02018 -0.1798 0.4208 1.0000
7.000 1.5693 0.03146 0.02125 -0.1774 0.4168 1.0000
7.250 1.5832 0.03221 0.02211 -0.1755 0.4130 1.0000
7.500 1.6007 0.03282 0.02276 -0.1741 0.4095 1.0000
7.750 1.6218 0.03330 0.02325 -0.1733 0.4065 1.0000
8.000 1.6465 0.03371 0.02361 -0.1730 0.4040 1.0000
8.250 1.6552 0.03465 0.02468 -0.1704 0.4007 1.0000
8.500 1.6561 0.03581 0.02601 -0.1668 0.3968 1.0000
8.750 1.6643 0.03676 0.02707 -0.1644 0.3932 1.0000
9.000 1.6796 0.03743 0.02777 -0.1628 0.3900 1.0000
9.250 1.7014 0.03784 0.02816 -0.1621 0.3873 1.0000
9.500 1.7113 0.03883 0.02923 -0.1601 0.3843 1.0000
9.750 1.6984 0.04099 0.03163 -0.1556 0.3800 1.0000
10.000 1.6980 0.04265 0.03344 -0.1528 0.3764 1.0000
10.250 1.7087 0.04372 0.03456 -0.1511 0.3733 1.0000
10.500 1.7291 0.04420 0.03505 -0.1504 0.3708 1.0000
10.750 1.7381 0.04543 0.03634 -0.1487 0.3678 1.0000
11.000 1.6941 0.05047 0.04172 -0.1432 0.3621 1.0000
11.250 1.6876 0.05307 0.04445 -0.1409 0.3580 1.0000
11.500 1.7050 0.05373 0.04514 -0.1401 0.3552 1.0000
11.750 1.7327 0.05357 0.04497 -0.1398 0.3532 1.0000
12.000 1.5842 0.07116 0.06309 -0.1339 0.3409 1.0000
12.250 1.6062 0.07128 0.06324 -0.1332 0.3390 1.0000
12.500 1.6377 0.07029 0.06226 -0.1326 0.3376 1.0000
13.000 1.5323 0.08996 0.08225 -0.1318 0.3208 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 441 AIRFOIL (goe441-il)