KENNEDY AND MARSDEN AIRFOIL (kenmar-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: KENNEDY AND MARSDEN AIRFOIL (kenmar-il) Reynolds number: 500,000 Max Cl/Cd: 28.26 at α=1.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-kenmar-il-500000.txt Download as CSV file: xf-kenmar-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: KENNEDY AND MARSDEN 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 0.4521 0.10671 0.10089 -0.1598 0.4559 0.0570
-10.500 0.4631 0.10483 0.09904 -0.1606 0.4556 0.0576
-10.250 0.2449 0.06254 0.05657 -0.1840 0.4572 0.0554
-10.000 0.2663 0.06201 0.05611 -0.1827 0.4568 0.0550
-9.750 0.1944 0.05093 0.04484 -0.1901 0.4570 0.0537
-9.500 0.1436 0.04462 0.03829 -0.1928 0.4569 0.0530
-9.250 0.1199 0.04120 0.03471 -0.1931 0.4567 0.0528
-9.000 0.1093 0.03896 0.03234 -0.1925 0.4565 0.0528
-8.750 0.1049 0.03728 0.03056 -0.1915 0.4563 0.0530
-8.500 0.1019 0.03591 0.02909 -0.1900 0.4561 0.0531
-8.250 0.0972 0.03493 0.02803 -0.1877 0.4559 0.0532
-8.000 0.0934 0.03410 0.02712 -0.1850 0.4557 0.0532
-7.750 0.0989 0.03321 0.02614 -0.1836 0.4554 0.0533
-7.500 0.1086 0.03242 0.02527 -0.1825 0.4551 0.0534
-7.250 0.1214 0.03173 0.02450 -0.1817 0.4548 0.0536
-7.000 0.1364 0.03109 0.02380 -0.1809 0.4545 0.0537
-6.750 0.1528 0.03055 0.02319 -0.1802 0.4542 0.0539
-6.500 0.1704 0.03008 0.02267 -0.1794 0.4539 0.0540
-6.250 0.1887 0.02969 0.02222 -0.1786 0.4535 0.0543
-6.000 0.2077 0.02934 0.02183 -0.1777 0.4532 0.0545
-5.750 0.2272 0.02905 0.02150 -0.1769 0.4528 0.0547
-5.500 0.2476 0.02876 0.02118 -0.1761 0.4523 0.0549
-5.250 0.2688 0.02850 0.02088 -0.1754 0.4517 0.0551
-5.000 0.2906 0.02830 0.02064 -0.1748 0.4511 0.0553
-4.750 0.3115 0.02809 0.02041 -0.1740 0.4507 0.0555
-4.500 0.3292 0.02774 0.02009 -0.1723 0.4503 0.0559
-4.250 0.3481 0.02753 0.01993 -0.1709 0.4499 0.0563
-4.000 0.3688 0.02740 0.01984 -0.1700 0.4495 0.0567
-3.750 0.3913 0.02728 0.01974 -0.1695 0.4490 0.0573
-3.500 0.4142 0.02719 0.01966 -0.1690 0.4484 0.0579
-3.250 0.4378 0.02713 0.01962 -0.1687 0.4479 0.0586
-3.000 0.4616 0.02708 0.01957 -0.1684 0.4474 0.0593
-2.750 0.4852 0.02708 0.01958 -0.1680 0.4469 0.0599
-2.500 0.5094 0.02711 0.01962 -0.1678 0.4465 0.0604
-2.250 0.5345 0.02717 0.01968 -0.1677 0.4461 0.0611
-2.000 0.5599 0.02723 0.01974 -0.1678 0.4457 0.0617
-1.750 0.5866 0.02717 0.01973 -0.1683 0.4453 0.0626
-1.500 0.6145 0.02727 0.01986 -0.1690 0.4448 0.0636
-1.250 0.6429 0.02746 0.02008 -0.1697 0.4443 0.0648
-1.000 0.6715 0.02773 0.02036 -0.1705 0.4438 0.0661
-0.750 0.6995 0.02815 0.02079 -0.1712 0.4433 0.0675
-0.500 0.7276 0.02881 0.02150 -0.1722 0.4426 0.0693
-0.250 0.7269 0.03131 0.02416 -0.1692 0.4407 0.0705
0.000 0.7545 0.03123 0.02412 -0.1697 0.4403 0.0731
0.250 0.7814 0.03127 0.02421 -0.1702 0.4398 0.0764
0.500 0.8053 0.03155 0.02458 -0.1704 0.4393 0.0813
0.750 0.8289 0.03194 0.02507 -0.1706 0.4387 0.0901
1.000 0.8544 0.03235 0.02573 -0.1717 0.4379 0.1251
1.250 0.8965 0.03279 0.02671 -0.1766 0.4368 0.2526
1.500 0.9306 0.03367 0.02799 -0.1800 0.4357 0.3517
1.750 0.9799 0.03467 0.02964 -0.1864 0.4345 0.5336
2.000 0.9663 0.03663 0.03179 -0.1816 0.4329 0.5594
2.250 0.9386 0.03932 0.03464 -0.1756 0.4310 0.5685
2.500 0.9204 0.04191 0.03735 -0.1717 0.4294 0.5819
2.750 0.9188 0.04380 0.03930 -0.1695 0.4282 0.5953
3.000 0.9341 0.04485 0.04033 -0.1689 0.4276 0.6105
3.250 0.9533 0.04535 0.04084 -0.1678 0.4271 0.6219
3.500 0.9693 0.04631 0.04177 -0.1673 0.4264 0.6315
3.750 0.9827 0.04717 0.04264 -0.1659 0.4258 0.6387
4.000 0.9944 0.04812 0.04360 -0.1643 0.4251 0.6464
4.250 1.0015 0.05003 0.04549 -0.1630 0.4242 0.6538
4.500 0.6582 0.09252 0.08857 -0.1643 0.3925 0.5962
4.750 0.6877 0.09227 0.08831 -0.1637 0.3930 0.6164
5.000 0.6310 0.10214 0.09825 -0.1647 0.3825 0.6100
5.250 0.6523 0.10258 0.09871 -0.1640 0.3821 0.6241
5.500 0.6766 0.10288 0.09896 -0.1637 0.3817 0.6356
5.750 0.6964 0.10320 0.09929 -0.1627 0.3813 0.6444
6.000 0.7184 0.10354 0.09959 -0.1619 0.3810 0.6542
6.250 0.7374 0.10393 0.09998 -0.1608 0.3807 0.6622
6.500 0.7512 0.10441 0.10049 -0.1587 0.3802 0.6700
6.750 0.7063 0.11312 0.10930 -0.1601 0.3693 0.6680
|
Polar data table (+)
Polar graphs
<< Back to KENNEDY AND MARSDEN AIRFOIL (kenmar-il)