USA-35B AIRFOIL (usa35b-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: USA-35B AIRFOIL (usa35b-il) Reynolds number: 50,000 Max Cl/Cd: 36.23 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-usa35b-il-50000-n5.txt Download as CSV file: xf-usa35b-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: USA-35B 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3126 0.10078 0.09366 -0.0322 1.0000 0.1485
-9.500 -0.3182 0.09862 0.09161 -0.0320 1.0000 0.1508
-9.000 -0.3682 0.08907 0.08226 -0.0379 1.0000 0.0928
-8.750 -0.3663 0.08663 0.07987 -0.0356 1.0000 0.0911
-8.500 -0.3713 0.08400 0.07732 -0.0347 1.0000 0.0896
-8.250 -0.3779 0.08116 0.07455 -0.0345 1.0000 0.0882
-7.750 -0.3897 0.06976 0.06286 -0.0440 0.9974 0.0805
-7.500 -0.3609 0.06531 0.05829 -0.0489 0.9888 0.0797
-7.250 -0.3311 0.06072 0.05351 -0.0543 0.9796 0.0786
-7.000 -0.2988 0.05587 0.04838 -0.0602 0.9706 0.0776
-6.750 -0.2674 0.05131 0.04347 -0.0652 0.9604 0.0768
-6.500 -0.2352 0.04721 0.03892 -0.0696 0.9501 0.0771
-6.250 -0.1975 0.04337 0.03450 -0.0743 0.9419 0.0786
-6.000 -0.1643 0.04031 0.03090 -0.0771 0.9311 0.0796
-5.750 -0.1286 0.03769 0.02776 -0.0797 0.9218 0.0801
-5.500 -0.0908 0.03538 0.02498 -0.0823 0.9130 0.0808
-5.250 -0.0570 0.03363 0.02304 -0.0840 0.9028 0.0822
-5.000 -0.0188 0.03221 0.02146 -0.0862 0.8948 0.0853
-4.750 0.0124 0.03106 0.02007 -0.0870 0.8832 0.0889
-4.500 0.0502 0.02974 0.01840 -0.0886 0.8751 0.0923
-4.250 0.0813 0.02865 0.01717 -0.0890 0.8638 0.0951
-4.000 0.1125 0.02778 0.01627 -0.0894 0.8534 0.0995
-3.750 0.1473 0.02698 0.01526 -0.0902 0.8446 0.1074
-3.500 0.1755 0.02633 0.01464 -0.0901 0.8329 0.1154
-3.250 0.2092 0.02559 0.01383 -0.0908 0.8245 0.1280
-3.000 0.2373 0.02507 0.01337 -0.0907 0.8132 0.1475
-2.750 0.2662 0.02451 0.01294 -0.0908 0.8028 0.1845
-2.500 0.2966 0.02347 0.01245 -0.0914 0.7940 0.2816
-2.250 0.3133 0.02198 0.01251 -0.0889 0.7834 0.6150
-2.000 0.3602 0.02114 0.01198 -0.0904 0.7757 1.0000
-1.750 0.3853 0.02139 0.01193 -0.0899 0.7636 1.0000
-1.500 0.4133 0.02153 0.01179 -0.0896 0.7534 1.0000
-1.250 0.4420 0.02163 0.01164 -0.0894 0.7434 1.0000
-1.000 0.4674 0.02186 0.01168 -0.0888 0.7317 1.0000
-0.750 0.4961 0.02196 0.01156 -0.0885 0.7221 1.0000
-0.500 0.5226 0.02216 0.01160 -0.0880 0.7112 1.0000
-0.250 0.5483 0.02242 0.01171 -0.0874 0.7004 1.0000
0.000 0.5779 0.02249 0.01162 -0.0872 0.6918 1.0000
0.250 0.6018 0.02287 0.01190 -0.0865 0.6804 1.0000
0.500 0.6295 0.02308 0.01199 -0.0861 0.6713 1.0000
0.750 0.6556 0.02336 0.01219 -0.0856 0.6612 1.0000
1.000 0.6810 0.02371 0.01247 -0.0851 0.6511 1.0000
1.250 0.7094 0.02391 0.01258 -0.0847 0.6422 1.0000
1.500 0.7329 0.02437 0.01301 -0.0840 0.6312 1.0000
1.750 0.7617 0.02456 0.01312 -0.0837 0.6226 1.0000
2.000 0.7847 0.02506 0.01363 -0.0829 0.6114 1.0000
2.250 0.8101 0.02544 0.01399 -0.0823 0.6013 1.0000
2.500 0.8365 0.02577 0.01430 -0.0817 0.5914 1.0000
2.750 0.8590 0.02632 0.01488 -0.0809 0.5801 1.0000
3.000 0.8871 0.02655 0.01505 -0.0804 0.5709 1.0000
3.250 0.9086 0.02713 0.01572 -0.0795 0.5588 1.0000
3.500 0.9317 0.02763 0.01625 -0.0786 0.5475 1.0000
3.750 0.9596 0.02780 0.01638 -0.0780 0.5374 1.0000
4.000 0.9795 0.02834 0.01702 -0.0767 0.5236 1.0000
4.250 1.0007 0.02878 0.01750 -0.0755 0.5099 1.0000
4.500 1.0228 0.02913 0.01788 -0.0743 0.4965 1.0000
4.750 1.0460 0.02939 0.01817 -0.0732 0.4837 1.0000
5.000 1.0687 0.02970 0.01851 -0.0721 0.4709 1.0000
5.250 1.0874 0.03028 0.01918 -0.0707 0.4572 1.0000
5.500 1.1067 0.03083 0.01984 -0.0693 0.4440 1.0000
5.750 1.1268 0.03133 0.02039 -0.0680 0.4313 1.0000
6.000 1.1482 0.03173 0.02083 -0.0668 0.4193 1.0000
6.250 1.1680 0.03224 0.02140 -0.0655 0.4067 1.0000
6.500 1.1847 0.03298 0.02226 -0.0639 0.3936 1.0000
6.750 1.2021 0.03369 0.02306 -0.0625 0.3811 1.0000
7.000 1.2207 0.03433 0.02374 -0.0610 0.3691 1.0000
7.250 1.2414 0.03484 0.02425 -0.0598 0.3577 1.0000
7.500 1.2547 0.03590 0.02546 -0.0580 0.3453 1.0000
7.750 1.2697 0.03691 0.02657 -0.0564 0.3342 1.0000
8.000 1.2892 0.03767 0.02733 -0.0552 0.3240 1.0000
8.250 1.3008 0.03888 0.02871 -0.0533 0.3129 1.0000
8.500 1.3109 0.04011 0.03004 -0.0514 0.3018 1.0000
8.750 1.3225 0.04103 0.03097 -0.0493 0.2907 1.0000
9.000 1.3266 0.04233 0.03236 -0.0468 0.2794 1.0000
9.250 1.3263 0.04401 0.03417 -0.0443 0.2684 1.0000
9.500 1.3285 0.04547 0.03567 -0.0420 0.2573 1.0000
9.750 1.3303 0.04697 0.03721 -0.0400 0.2462 1.0000
10.000 1.3232 0.04954 0.03996 -0.0382 0.2359 1.0000
10.250 1.3213 0.05174 0.04223 -0.0368 0.2259 1.0000
10.500 1.3211 0.05384 0.04436 -0.0356 0.2160 1.0000
10.750 1.3134 0.05712 0.04784 -0.0349 0.2068 1.0000
11.000 1.3123 0.05966 0.05042 -0.0342 0.1980 1.0000
11.250 1.3071 0.06289 0.05380 -0.0339 0.1891 1.0000
11.500 1.3011 0.06638 0.05742 -0.0338 0.1805 1.0000
11.750 1.3005 0.06899 0.05998 -0.0335 0.1714 1.0000
12.000 1.2875 0.07380 0.06504 -0.0342 0.1628 1.0000
12.250 1.2817 0.07744 0.06872 -0.0345 0.1540 1.0000
12.500 1.2739 0.08152 0.07287 -0.0350 0.1452 1.0000
12.750 1.2616 0.08659 0.07808 -0.0361 0.1371 1.0000
13.000 1.2600 0.08969 0.08108 -0.0363 0.1288 1.0000
13.250 1.2417 0.09632 0.08798 -0.0385 0.1220 1.0000
13.500 1.2427 0.09915 0.09070 -0.0387 0.1151 1.0000
13.750 1.2221 0.10668 0.09852 -0.0417 0.1100 1.0000
14.000 1.2174 0.11101 0.10289 -0.0429 0.1048 1.0000
14.250 1.2078 0.11655 0.10850 -0.0450 0.1007 1.0000
14.500 1.1774 0.12715 0.11936 -0.0504 0.0976 1.0000
14.750 1.1562 0.13609 0.12842 -0.0550 0.0943 1.0000
15.000 1.1705 0.13643 0.12870 -0.0543 0.0904 1.0000
15.250 1.0972 0.16080 0.15320 -0.0688 0.0880 1.0000
15.500 1.0750 0.17290 0.16525 -0.0757 0.0846 1.0000
|
Polar data table (+)
Polar graphs
<< Back to USA-35B AIRFOIL (usa35b-il)