FX 66-S-196 AIRFOIL (fx66s196-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-S-196 AIRFOIL (fx66s196-il) Reynolds number: 200,000 Max Cl/Cd: 37.19 at α=10.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx66s196-il-200000.txt Download as CSV file: xf-fx66s196-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-S-196 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-13.000 -0.1543 0.10177 0.09779 -0.0947 0.7728 0.0748
-12.750 -0.1291 0.09933 0.09527 -0.0921 0.7666 0.0767
-12.500 -0.3803 0.06579 0.06128 -0.1232 0.8073 0.0421
-12.250 -0.3747 0.06277 0.05812 -0.1233 0.7995 0.0408
-12.000 -0.3920 0.05834 0.05358 -0.1240 0.7912 0.0402
-11.750 -0.4106 0.05456 0.04959 -0.1241 0.7847 0.0396
-11.500 -0.4289 0.05119 0.04606 -0.1237 0.7770 0.0392
-11.250 -0.4464 0.04780 0.04242 -0.1225 0.7708 0.0383
-11.000 -0.4642 0.04469 0.03906 -0.1206 0.7643 0.0376
-10.750 -0.4818 0.04183 0.03587 -0.1176 0.7583 0.0368
-10.500 -0.4926 0.03906 0.03268 -0.1144 0.7537 0.0361
-10.250 -0.4918 0.03663 0.02994 -0.1124 0.7478 0.0359
-10.000 -0.4819 0.03465 0.02769 -0.1110 0.7426 0.0362
-9.750 -0.4674 0.03301 0.02580 -0.1099 0.7384 0.0368
-9.500 -0.4516 0.03173 0.02438 -0.1090 0.7330 0.0380
-9.250 -0.4341 0.03044 0.02287 -0.1081 0.7285 0.0392
-9.000 -0.4134 0.02901 0.02117 -0.1074 0.7248 0.0402
-8.750 -0.3912 0.02785 0.01979 -0.1068 0.7211 0.0410
-8.500 -0.3660 0.02646 0.01831 -0.1065 0.7168 0.0419
-8.250 -0.3397 0.02502 0.01692 -0.1065 0.7131 0.0440
-8.000 -0.3191 0.02432 0.01618 -0.1058 0.7098 0.0466
-7.750 -0.2983 0.02374 0.01546 -0.1050 0.7070 0.0496
-7.500 -0.2804 0.02283 0.01458 -0.1038 0.7032 0.0525
-7.250 -0.2659 0.02201 0.01383 -0.1024 0.6995 0.0566
-7.000 -0.2487 0.02141 0.01315 -0.1012 0.6961 0.0619
-6.750 -0.2351 0.02058 0.01234 -0.0998 0.6934 0.0696
-6.500 -0.2196 0.01989 0.01165 -0.0986 0.6908 0.0813
-6.250 -0.2049 0.01915 0.01102 -0.0974 0.6874 0.1004
-6.000 -0.1922 0.01807 0.01027 -0.0962 0.6842 0.1472
-5.750 -0.1835 0.01645 0.00939 -0.0949 0.6813 0.2895
-5.500 -0.1686 0.01603 0.00992 -0.0931 0.6788 0.4882
-5.250 -0.1408 0.01653 0.01018 -0.0929 0.6766 0.5306
-5.000 -0.1142 0.01716 0.01079 -0.0923 0.6741 0.5502
-4.750 -0.0880 0.01774 0.01133 -0.0918 0.6712 0.5654
-4.500 -0.0613 0.01829 0.01182 -0.0912 0.6683 0.5795
-4.250 -0.0344 0.01880 0.01225 -0.0906 0.6658 0.5930
-4.000 -0.0065 0.01930 0.01272 -0.0901 0.6637 0.6026
-3.750 0.0211 0.01962 0.01294 -0.0897 0.6619 0.6111
-3.500 0.0483 0.02004 0.01330 -0.0894 0.6602 0.6197
-3.250 0.0722 0.02055 0.01386 -0.0883 0.6577 0.6284
-3.000 0.0967 0.02094 0.01420 -0.0879 0.6553 0.6398
-2.750 0.1218 0.02138 0.01469 -0.0867 0.6531 0.6450
-2.500 0.1478 0.02153 0.01478 -0.0866 0.6511 0.6513
-2.250 0.1739 0.02159 0.01475 -0.0866 0.6493 0.6578
-2.000 0.2004 0.02180 0.01493 -0.0862 0.6477 0.6621
-1.750 0.2281 0.02182 0.01485 -0.0866 0.6463 0.6665
-1.500 0.2555 0.02181 0.01471 -0.0875 0.6447 0.6717
-1.250 0.2761 0.02204 0.01502 -0.0866 0.6419 0.6741
-1.000 0.2992 0.02224 0.01524 -0.0862 0.6393 0.6766
-0.750 0.3240 0.02237 0.01534 -0.0863 0.6372 0.6790
-0.500 0.3503 0.02243 0.01536 -0.0867 0.6354 0.6810
-0.250 0.3773 0.02246 0.01533 -0.0873 0.6337 0.6829
0.000 0.4052 0.02249 0.01527 -0.0880 0.6323 0.6848
0.250 0.4339 0.02254 0.01523 -0.0889 0.6310 0.6869
0.500 0.4623 0.02263 0.01527 -0.0896 0.6298 0.6884
0.750 0.4792 0.02321 0.01593 -0.0886 0.6272 0.6895
1.000 0.4943 0.02388 0.01668 -0.0875 0.6240 0.6907
1.250 0.5152 0.02427 0.01710 -0.0872 0.6213 0.6919
1.500 0.5389 0.02456 0.01740 -0.0871 0.6193 0.6933
1.750 0.5646 0.02479 0.01761 -0.0874 0.6178 0.6947
2.000 0.5925 0.02494 0.01774 -0.0879 0.6166 0.6964
2.250 0.6216 0.02504 0.01781 -0.0885 0.6153 0.6984
2.500 0.6515 0.02517 0.01789 -0.0894 0.6142 0.7001
2.750 0.6167 0.02836 0.02133 -0.0822 0.6055 0.7014
3.000 0.6352 0.02898 0.02195 -0.0817 0.6032 0.7026
3.250 0.6603 0.02929 0.02226 -0.0818 0.6018 0.7043
3.500 0.6886 0.02941 0.02241 -0.0822 0.6007 0.7061
3.750 0.7201 0.02943 0.02243 -0.0831 0.5999 0.7076
4.000 0.7541 0.02936 0.02237 -0.0843 0.5991 0.7094
4.250 0.5144 0.04506 0.03825 -0.0588 0.5701 0.7105
4.500 0.5500 0.04473 0.03793 -0.0596 0.5691 0.7123
4.750 0.5664 0.04571 0.03892 -0.0590 0.5657 0.7139
5.000 0.6146 0.04456 0.03778 -0.0608 0.5666 0.7159
5.250 0.6584 0.04371 0.03695 -0.0623 0.5672 0.7179
5.500 0.5992 0.04963 0.04289 -0.0568 0.5532 0.7197
5.750 0.6357 0.04921 0.04251 -0.0575 0.5525 0.7216
6.000 0.6703 0.04900 0.04236 -0.0582 0.5519 0.7239
7.500 0.6922 0.05978 0.05345 -0.0520 0.5114 0.7417
7.750 0.7227 0.05975 0.05350 -0.0525 0.5094 0.7465
8.000 0.7560 0.05947 0.05329 -0.0531 0.5081 0.7519
8.250 0.7448 0.06260 0.05651 -0.0517 0.4972 0.7561
8.500 0.7741 0.06250 0.05651 -0.0521 0.4947 0.7630
8.750 0.8101 0.06179 0.05591 -0.0526 0.4932 0.7709
9.000 0.8479 0.06091 0.05515 -0.0533 0.4921 0.7807
9.250 0.8870 0.05992 0.05433 -0.0541 0.4913 0.7928
9.500 0.9296 0.05871 0.05330 -0.0552 0.4908 0.8093
10.500 1.2854 0.03456 0.02999 -0.0690 0.5053 1.0000
10.750 1.0640 0.05808 0.05343 -0.0590 0.4633 1.0000
11.000 1.1172 0.05460 0.05003 -0.0596 0.4629 1.0000
11.250 1.1821 0.04942 0.04497 -0.0604 0.4632 1.0000
11.500 1.2718 0.04165 0.03732 -0.0621 0.4634 1.0000
11.750 1.2679 0.04391 0.03964 -0.0609 0.4503 1.0000
12.000 1.2787 0.04481 0.04060 -0.0603 0.4372 1.0000
12.250 1.3035 0.04427 0.04010 -0.0601 0.4192 1.0000
12.500 1.3078 0.04600 0.04184 -0.0592 0.3966 1.0000
13.000 1.3402 0.04595 0.04103 -0.0568 0.3024 1.0000
13.250 1.3224 0.04990 0.04473 -0.0551 0.2664 1.0000
13.500 1.3035 0.05421 0.04881 -0.0537 0.2334 1.0000
13.750 1.2836 0.05883 0.05319 -0.0526 0.1997 1.0000
14.000 1.2651 0.06348 0.05761 -0.0516 0.1656 1.0000
14.250 1.2470 0.06825 0.06213 -0.0509 0.1323 1.0000
14.500 1.2300 0.07298 0.06661 -0.0504 0.1008 1.0000
14.750 1.2137 0.07777 0.07114 -0.0499 0.0742 1.0000
15.000 1.2024 0.08213 0.07537 -0.0496 0.0587 1.0000
15.250 1.1946 0.08615 0.07932 -0.0495 0.0484 1.0000
15.500 1.1899 0.08989 0.08305 -0.0494 0.0428 1.0000
15.750 1.1856 0.09359 0.08679 -0.0494 0.0383 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-S-196 AIRFOIL (fx66s196-il)