BELL 540 AIRFOIL (MODIFIED NACA 0012) (b540ols-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BELL 540 AIRFOIL (MODIFIED NACA 0012) (b540ols-il) Reynolds number: 50,000 Max Cl/Cd: 24.9 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-b540ols-il-50000-n5.txt Download as CSV file: xf-b540ols-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BELL 540 AIRFOIL (MODIFIED NACA 0012)
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
------ -------- --------- --------- -------- -------- --------
-10.250 -0.8099 0.08663 0.07907 0.0087 1.0000 0.0736
-10.000 -0.8631 0.07175 0.06404 -0.0046 1.0000 0.0715
-9.750 -0.8927 0.06407 0.05607 -0.0087 1.0000 0.0713
-9.500 -0.9028 0.05865 0.05031 -0.0103 1.0000 0.0725
-9.250 -0.9059 0.05372 0.04494 -0.0112 1.0000 0.0743
-9.000 -0.9035 0.04903 0.03965 -0.0115 1.0000 0.0764
-8.750 -0.8926 0.04533 0.03549 -0.0114 1.0000 0.0784
-8.500 -0.8741 0.04337 0.03350 -0.0110 1.0000 0.0814
-8.250 -0.8567 0.04102 0.03085 -0.0107 1.0000 0.0856
-8.000 -0.8387 0.03818 0.02744 -0.0103 1.0000 0.0899
-7.750 -0.8172 0.03656 0.02589 -0.0098 1.0000 0.0940
-7.500 -0.7957 0.03471 0.02377 -0.0094 1.0000 0.1002
-7.250 -0.7734 0.03299 0.02192 -0.0089 1.0000 0.1060
-7.000 -0.7504 0.03149 0.02025 -0.0085 1.0000 0.1138
-6.750 -0.7272 0.03009 0.01883 -0.0080 1.0000 0.1218
-6.500 -0.7033 0.02872 0.01731 -0.0075 1.0000 0.1321
-6.250 -0.6792 0.02751 0.01598 -0.0071 1.0000 0.1444
-6.000 -0.6551 0.02639 0.01488 -0.0066 1.0000 0.1585
-5.750 -0.6307 0.02533 0.01385 -0.0062 1.0000 0.1752
-5.500 -0.6063 0.02439 0.01299 -0.0058 1.0000 0.1946
-5.250 -0.5813 0.02352 0.01215 -0.0054 1.0000 0.2184
-5.000 -0.5563 0.02276 0.01151 -0.0050 1.0000 0.2438
-4.750 -0.5310 0.02211 0.01093 -0.0046 1.0000 0.2725
-4.500 -0.5054 0.02156 0.01043 -0.0041 1.0000 0.3024
-4.250 -0.4797 0.02108 0.01000 -0.0036 1.0000 0.3325
-4.000 -0.4537 0.02065 0.00959 -0.0032 1.0000 0.3627
-3.750 -0.4282 0.02024 0.00922 -0.0026 1.0000 0.3912
-3.500 -0.4024 0.01983 0.00886 -0.0020 1.0000 0.4189
-3.250 -0.3765 0.01941 0.00848 -0.0014 1.0000 0.4467
-3.000 -0.3506 0.01900 0.00812 -0.0008 1.0000 0.4738
-2.750 -0.3249 0.01859 0.00781 -0.0001 1.0000 0.5005
-2.500 -0.2990 0.01819 0.00750 0.0005 1.0000 0.5285
-2.250 -0.2735 0.01779 0.00722 0.0013 1.0000 0.5565
-2.000 -0.2483 0.01739 0.00699 0.0021 1.0000 0.5862
-1.750 -0.2234 0.01700 0.00679 0.0031 1.0000 0.6186
-1.500 -0.1990 0.01662 0.00662 0.0042 1.0000 0.6552
-1.250 -0.1752 0.01625 0.00651 0.0056 1.0000 0.6976
-1.000 -0.1514 0.01591 0.00646 0.0072 1.0000 0.7495
-0.750 -0.1241 0.01562 0.00644 0.0082 1.0000 0.8163
-0.500 -0.0827 0.01542 0.00643 0.0063 1.0000 0.8993
-0.250 -0.0288 0.01528 0.00635 0.0011 1.0000 0.9822
0.000 0.0000 0.01522 0.00629 0.0000 1.0000 1.0000
0.250 0.0288 0.01528 0.00635 -0.0011 0.9822 1.0000
0.500 0.0827 0.01542 0.00643 -0.0063 0.8993 1.0000
0.750 0.1241 0.01562 0.00644 -0.0082 0.8163 1.0000
1.000 0.1514 0.01591 0.00646 -0.0072 0.7495 1.0000
1.250 0.1752 0.01625 0.00651 -0.0056 0.6976 1.0000
1.500 0.1989 0.01662 0.00662 -0.0042 0.6552 1.0000
1.750 0.2234 0.01700 0.00679 -0.0031 0.6186 1.0000
2.000 0.2482 0.01739 0.00699 -0.0021 0.5862 1.0000
2.250 0.2735 0.01779 0.00722 -0.0013 0.5565 1.0000
2.500 0.2990 0.01819 0.00750 -0.0005 0.5285 1.0000
2.750 0.3248 0.01859 0.00781 0.0001 0.5005 1.0000
3.000 0.3506 0.01900 0.00812 0.0008 0.4739 1.0000
3.250 0.3765 0.01941 0.00848 0.0014 0.4467 1.0000
3.500 0.4024 0.01983 0.00886 0.0020 0.4189 1.0000
3.750 0.4282 0.02024 0.00922 0.0026 0.3912 1.0000
4.000 0.4537 0.02065 0.00959 0.0032 0.3627 1.0000
4.250 0.4796 0.02108 0.01000 0.0036 0.3325 1.0000
4.500 0.5054 0.02155 0.01043 0.0041 0.3024 1.0000
4.750 0.5309 0.02211 0.01093 0.0046 0.2725 1.0000
5.000 0.5563 0.02276 0.01151 0.0050 0.2438 1.0000
5.250 0.5813 0.02352 0.01215 0.0054 0.2184 1.0000
5.500 0.6063 0.02439 0.01299 0.0058 0.1946 1.0000
5.750 0.6307 0.02533 0.01385 0.0062 0.1752 1.0000
6.000 0.6551 0.02639 0.01488 0.0066 0.1585 1.0000
6.250 0.6792 0.02751 0.01598 0.0071 0.1444 1.0000
6.500 0.7034 0.02872 0.01731 0.0075 0.1321 1.0000
6.750 0.7273 0.03009 0.01883 0.0080 0.1218 1.0000
7.000 0.7504 0.03149 0.02025 0.0085 0.1138 1.0000
7.250 0.7734 0.03299 0.02194 0.0089 0.1060 1.0000
7.500 0.7957 0.03471 0.02377 0.0094 0.1002 1.0000
7.750 0.8173 0.03656 0.02589 0.0098 0.0940 1.0000
8.000 0.8388 0.03818 0.02744 0.0103 0.0899 1.0000
8.250 0.8567 0.04102 0.03085 0.0106 0.0856 1.0000
8.500 0.8742 0.04337 0.03351 0.0110 0.0814 1.0000
8.750 0.8927 0.04533 0.03550 0.0114 0.0784 1.0000
9.000 0.9036 0.04904 0.03966 0.0115 0.0764 1.0000
9.250 0.9060 0.05374 0.04495 0.0111 0.0743 1.0000
9.500 0.9030 0.05866 0.05032 0.0103 0.0725 1.0000
9.750 0.8930 0.06408 0.05608 0.0086 0.0713 1.0000
10.000 0.8634 0.07177 0.06407 0.0045 0.0715 1.0000
10.250 0.8101 0.08676 0.07919 -0.0089 0.0735 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BELL 540 AIRFOIL (MODIFIED NACA 0012) (b540ols-il)