FX 66-S-161 AIRFOIL (fx66s161-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: FX 66-S-161 AIRFOIL (fx66s161-il) Reynolds number: 50,000 Max Cl/Cd: 5.32 at α=9° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-fx66s161-il-50000.txt Download as CSV file: xf-fx66s161-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: FX 66-S-161 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.2431 0.11589 0.11002 -0.0367 1.0000 0.2865
-9.750 -0.2420 0.11325 0.10753 -0.0363 1.0000 0.2946
-9.500 -0.2583 0.11317 0.10769 -0.0351 1.0000 0.3044
-9.250 -0.2967 0.11645 0.11132 -0.0271 1.0000 0.3038
-9.000 -0.3388 0.11987 0.11496 -0.0209 0.9999 0.3044
-8.750 -0.3122 0.11488 0.10993 -0.0264 0.9906 0.3196
-8.500 -0.2773 0.10914 0.10410 -0.0315 0.9822 0.3316
-8.250 -0.2640 0.07782 0.07309 -0.0720 0.9373 0.1583
-8.000 -0.2559 0.07196 0.06719 -0.0751 0.9302 0.1507
-7.750 -0.4065 0.07508 0.07008 -0.0760 0.9430 0.1397
-7.500 -0.4454 0.06741 0.06192 -0.0825 0.9299 0.1286
-7.250 -0.4326 0.06308 0.05741 -0.0842 0.9214 0.1246
-7.000 -0.4240 0.05798 0.05187 -0.0871 0.9136 0.1201
-6.750 -0.4243 0.05405 0.04728 -0.0875 0.9054 0.1162
-6.500 -0.4038 0.05037 0.04298 -0.0893 0.8987 0.1160
-6.250 -0.3874 0.04821 0.04058 -0.0892 0.8917 0.1196
-6.000 -0.3658 0.04584 0.03773 -0.0897 0.8851 0.1226
-5.750 -0.3359 0.04341 0.03467 -0.0908 0.8796 0.1261
-5.500 -0.3246 0.04216 0.03328 -0.0891 0.8735 0.1311
-5.250 -0.2941 0.04068 0.03148 -0.0898 0.8685 0.1415
-5.000 -0.2713 0.03943 0.03028 -0.0892 0.8635 0.1525
-4.750 -0.2581 0.03872 0.02957 -0.0874 0.8581 0.1665
-4.500 -0.2312 0.03745 0.02846 -0.0872 0.8535 0.1928
-4.250 -0.2116 0.03606 0.02757 -0.0867 0.8494 0.2483
-4.000 -0.2137 0.03584 0.02893 -0.0818 0.8455 0.4263
-3.750 -0.2143 0.03849 0.03173 -0.0740 0.8414 0.5370
-3.500 -0.2031 0.04030 0.03343 -0.0687 0.8376 0.5909
-3.250 -0.1992 0.04160 0.03469 -0.0629 0.8345 0.6242
-3.000 -0.2010 0.04250 0.03554 -0.0575 0.8320 0.6527
-2.750 -0.2006 0.04327 0.03626 -0.0522 0.8298 0.6818
-2.500 -0.2010 0.04392 0.03689 -0.0464 0.8284 0.7088
-2.250 -0.2011 0.04439 0.03732 -0.0412 0.8278 0.7354
-2.000 -0.1983 0.04470 0.03753 -0.0372 0.8274 0.7605
-1.750 -0.1944 0.04493 0.03767 -0.0336 0.8271 0.7823
-1.500 -0.1886 0.04513 0.03774 -0.0313 0.8282 0.8024
-1.250 -0.1814 0.04538 0.03786 -0.0294 0.8311 0.8197
-1.000 -0.1692 0.04570 0.03804 -0.0284 0.8335 0.8361
-0.750 -0.1726 0.04589 0.03819 -0.0262 0.8446 0.8501
-0.500 -0.1514 0.04666 0.03880 -0.0271 0.8510 0.8663
0.000 -0.2932 0.04176 0.03446 -0.0013 1.0000 0.8842
0.250 -0.2695 0.04225 0.03482 -0.0028 1.0000 0.9061
0.500 -0.2376 0.04300 0.03546 -0.0064 1.0000 0.9314
0.750 -0.1989 0.04386 0.03619 -0.0122 0.9987 0.9627
1.000 -0.1509 0.04573 0.03781 -0.0204 0.9871 1.0000
1.250 -0.1100 0.04767 0.03947 -0.0271 0.9768 1.0000
1.500 -0.0685 0.04981 0.04132 -0.0335 0.9665 1.0000
1.750 -0.0225 0.05255 0.04377 -0.0406 0.9566 1.0000
2.000 0.0120 0.05418 0.04517 -0.0452 0.9435 1.0000
2.250 0.0445 0.05588 0.04666 -0.0491 0.9302 1.0000
2.500 0.0751 0.05767 0.04824 -0.0525 0.9173 1.0000
2.750 0.1048 0.05965 0.05002 -0.0556 0.9066 1.0000
3.000 0.1442 0.06261 0.05277 -0.0601 0.8966 1.0000
3.250 0.1672 0.06395 0.05396 -0.0616 0.8828 1.0000
3.500 0.1885 0.06540 0.05527 -0.0627 0.8698 1.0000
3.750 0.2097 0.06707 0.05682 -0.0638 0.8582 1.0000
4.000 0.2419 0.06988 0.05950 -0.0666 0.8501 1.0000
4.250 0.2651 0.07158 0.06111 -0.0678 0.8373 1.0000
4.500 0.2799 0.07283 0.06229 -0.0677 0.8250 1.0000
4.750 0.2977 0.07457 0.06396 -0.0682 0.8147 1.0000
5.000 0.3326 0.07776 0.06706 -0.0711 0.8065 1.0000
5.250 0.3432 0.07870 0.06798 -0.0704 0.7939 1.0000
5.500 0.3564 0.08029 0.06953 -0.0703 0.7839 1.0000
5.750 0.3899 0.08353 0.07273 -0.0729 0.7765 1.0000
6.000 0.3975 0.08446 0.07365 -0.0719 0.7643 1.0000
6.250 0.4111 0.08630 0.07549 -0.0719 0.7551 1.0000
6.500 0.4398 0.08918 0.07835 -0.0737 0.7468 1.0000
6.750 0.4461 0.09035 0.07954 -0.0728 0.7357 1.0000
7.000 0.4730 0.09357 0.08276 -0.0745 0.7294 1.0000
7.250 0.4838 0.09493 0.08416 -0.0742 0.7179 1.0000
7.500 0.4926 0.09668 0.08594 -0.0738 0.7087 1.0000
7.750 0.5225 0.10000 0.08929 -0.0757 0.7009 1.0000
8.000 0.5243 0.10108 0.09041 -0.0746 0.6902 1.0000
8.250 0.5545 0.10495 0.09434 -0.0766 0.6843 1.0000
8.500 0.5580 0.10593 0.09539 -0.0756 0.6726 1.0000
8.750 0.5658 0.10801 0.09751 -0.0755 0.6643 1.0000
9.000 0.5921 0.11124 0.10082 -0.0769 0.6559 1.0000
9.250 0.5930 0.11264 0.10228 -0.0761 0.6457 1.0000
9.750 0.6228 0.11779 0.10762 -0.0772 0.6280 1.0000
10.000 0.6314 0.12025 0.11016 -0.0773 0.6197 1.0000
10.250 0.6559 0.12364 0.11366 -0.0785 0.6107 1.0000
10.500 0.6552 0.12520 0.11530 -0.0781 0.6004 1.0000
10.750 0.6816 0.12946 0.11970 -0.0796 0.5939 1.0000
11.000 0.6845 0.13093 0.12127 -0.0793 0.5822 1.0000
11.250 0.6881 0.13325 0.12368 -0.0795 0.5732 1.0000
11.500 0.7186 0.13779 0.12838 -0.0811 0.5649 1.0000
|
Polar data table (+)
Polar graphs
<< Back to FX 66-S-161 AIRFOIL (fx66s161-il)