USA 41 AIRFOIL (usa41-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA 41 AIRFOIL (usa41-il) Reynolds number: 50,000 Max Cl/Cd: 42.19 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa41-il-50000-n5.txt Download as CSV file: xf-usa41-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA 41 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-7.750 -0.3309 0.10094 0.09423 -0.0252 1.0000 0.1024
-7.500 -0.3386 0.09993 0.09338 -0.0255 1.0000 0.1048
-7.250 -0.3464 0.09933 0.09292 -0.0287 1.0000 0.1060
-7.000 -0.3347 0.09401 0.08764 -0.0250 1.0000 0.1095
-6.750 -0.3306 0.09131 0.08500 -0.0243 1.0000 0.1136
-6.500 -0.3308 0.08932 0.08311 -0.0257 1.0000 0.1178
-6.250 -0.3315 0.08824 0.08212 -0.0316 1.0000 0.1203
-6.000 -0.3268 0.08464 0.07860 -0.0312 1.0000 0.1212
-5.750 -0.3220 0.08109 0.07511 -0.0287 1.0000 0.1229
-5.500 -0.3167 0.07817 0.07223 -0.0283 1.0000 0.1245
-5.250 -0.3103 0.07533 0.06942 -0.0288 1.0000 0.1265
-5.000 -0.2789 0.06752 0.06131 -0.0435 1.0000 0.0783
-4.750 -0.2692 0.06433 0.05811 -0.0427 1.0000 0.0765
-4.500 -0.2563 0.06089 0.05463 -0.0434 1.0000 0.0749
-4.250 -0.2385 0.05698 0.05063 -0.0457 1.0000 0.0713
-3.750 -0.1844 0.04773 0.04083 -0.0545 1.0000 0.0669
-3.500 -0.1626 0.04510 0.03801 -0.0559 1.0000 0.0702
-3.250 -0.1295 0.04117 0.03369 -0.0597 0.9984 0.0703
-3.000 -0.0820 0.03663 0.02853 -0.0658 0.9921 0.0696
-2.750 -0.0376 0.03338 0.02467 -0.0704 0.9852 0.0723
-2.500 0.0062 0.03050 0.02099 -0.0743 0.9779 0.0763
-2.250 0.0468 0.02819 0.01805 -0.0770 0.9697 0.0769
-2.000 0.0872 0.02644 0.01579 -0.0795 0.9613 0.0781
-1.750 0.1275 0.02520 0.01428 -0.0820 0.9527 0.0807
-1.500 0.1648 0.02427 0.01310 -0.0837 0.9420 0.0843
-1.250 0.2051 0.02351 0.01199 -0.0859 0.9319 0.0912
-0.750 0.2883 0.02237 0.01084 -0.0909 0.9114 0.1261
-0.500 0.3277 0.02191 0.01039 -0.0928 0.8994 0.1838
-0.250 0.3668 0.02130 0.00992 -0.0951 0.8875 0.2664
0.000 0.4042 0.02048 0.00966 -0.0970 0.8751 0.3721
0.500 0.4753 0.01904 0.00906 -0.0987 0.8464 1.0000
0.750 0.5076 0.01909 0.00888 -0.0991 0.8303 1.0000
1.000 0.5392 0.01913 0.00872 -0.0992 0.8138 1.0000
1.250 0.5704 0.01916 0.00858 -0.0993 0.7974 1.0000
1.500 0.6012 0.01919 0.00847 -0.0992 0.7810 1.0000
1.750 0.6291 0.01930 0.00848 -0.0987 0.7628 1.0000
2.000 0.6573 0.01942 0.00849 -0.0983 0.7449 1.0000
2.250 0.6859 0.01954 0.00852 -0.0978 0.7275 1.0000
2.500 0.7143 0.01969 0.00859 -0.0974 0.7104 1.0000
2.750 0.7425 0.01987 0.00871 -0.0969 0.6936 1.0000
3.000 0.7693 0.02012 0.00891 -0.0963 0.6762 1.0000
3.250 0.7955 0.02041 0.00918 -0.0957 0.6587 1.0000
3.500 0.8217 0.02073 0.00952 -0.0950 0.6419 1.0000
3.750 0.8478 0.02107 0.00986 -0.0944 0.6257 1.0000
4.000 0.8738 0.02144 0.01024 -0.0938 0.6099 1.0000
4.250 0.8996 0.02184 0.01071 -0.0931 0.5945 1.0000
4.500 0.9252 0.02226 0.01120 -0.0925 0.5794 1.0000
4.750 0.9506 0.02273 0.01173 -0.0918 0.5646 1.0000
5.000 0.9757 0.02322 0.01232 -0.0911 0.5500 1.0000
5.250 1.0004 0.02375 0.01301 -0.0903 0.5356 1.0000
5.500 1.0249 0.02430 0.01371 -0.0896 0.5213 1.0000
5.750 1.0490 0.02489 0.01446 -0.0888 0.5072 1.0000
6.000 1.0729 0.02550 0.01526 -0.0879 0.4933 1.0000
6.250 1.0964 0.02614 0.01617 -0.0870 0.4794 1.0000
6.500 1.1197 0.02681 0.01708 -0.0860 0.4657 1.0000
6.750 1.1424 0.02747 0.01802 -0.0850 0.4514 1.0000
7.000 1.1632 0.02792 0.01870 -0.0833 0.4320 1.0000
7.250 1.1729 0.02780 0.01857 -0.0796 0.3891 1.0000
7.500 1.1761 0.02805 0.01865 -0.0754 0.3276 1.0000
7.750 1.1816 0.02901 0.01944 -0.0721 0.2571 1.0000
8.000 1.1794 0.03130 0.02093 -0.0686 0.1422 1.0000
8.250 1.1738 0.03470 0.02361 -0.0654 0.0839 1.0000
8.500 1.1733 0.03739 0.02624 -0.0625 0.0684 1.0000
8.750 1.1712 0.03985 0.02872 -0.0594 0.0609 1.0000
9.000 1.1711 0.04212 0.03114 -0.0568 0.0546 1.0000
9.250 1.1673 0.04477 0.03385 -0.0544 0.0506 1.0000
9.500 1.1681 0.04722 0.03649 -0.0524 0.0473 1.0000
9.750 1.1695 0.04978 0.03924 -0.0506 0.0450 1.0000
10.000 1.1722 0.05236 0.04195 -0.0490 0.0428 1.0000
10.250 1.1754 0.05505 0.04469 -0.0476 0.0405 1.0000
10.500 1.1865 0.05742 0.04728 -0.0459 0.0378 1.0000
10.750 1.2003 0.05980 0.05002 -0.0443 0.0355 1.0000
11.000 1.2163 0.06244 0.05292 -0.0428 0.0341 1.0000
11.250 1.2279 0.06557 0.05632 -0.0416 0.0330 1.0000
11.500 1.2331 0.06900 0.05999 -0.0407 0.0321 1.0000
11.750 1.2356 0.07280 0.06395 -0.0401 0.0311 1.0000
12.000 1.2333 0.07715 0.06850 -0.0398 0.0304 1.0000
12.250 1.2220 0.08170 0.07338 -0.0402 0.0302 1.0000
12.500 1.2092 0.08669 0.07867 -0.0413 0.0300 1.0000
12.750 1.1950 0.09215 0.08440 -0.0431 0.0299 1.0000
13.000 1.1805 0.09802 0.09051 -0.0455 0.0300 1.0000
13.250 1.1650 0.10439 0.09710 -0.0486 0.0300 1.0000
13.500 1.1495 0.11118 0.10407 -0.0523 0.0301 1.0000
13.750 1.1344 0.11831 0.11135 -0.0564 0.0303 1.0000
14.000 1.1197 0.12583 0.11900 -0.0609 0.0304 1.0000
14.250 1.0744 0.14410 0.13765 -0.0745 0.0338 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA 41 AIRFOIL (usa41-il)