ONERA HOR07 AIRFOIL (hor07-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: ONERA HOR07 AIRFOIL (hor07-il) Reynolds number: 50,000 Max Cl/Cd: 43.41 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-hor07-il-50000-n5.txt Download as CSV file: xf-hor07-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: ONERA HOR07 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
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4062 0.11167 0.10486 -0.0056 1.0000 0.0992
-8.000 -0.4056 0.10955 0.10282 -0.0085 1.0000 0.1016
-7.750 -0.4058 0.10754 0.10090 -0.0121 1.0000 0.1024
-7.500 -0.3995 0.10472 0.09815 -0.0170 1.0000 0.1028
-7.250 -0.3881 0.10023 0.09371 -0.0174 1.0000 0.1034
-7.000 -0.3763 0.09610 0.08962 -0.0177 1.0000 0.1042
-6.750 -0.3648 0.09241 0.08597 -0.0191 1.0000 0.1052
-6.500 -0.3527 0.08891 0.08250 -0.0210 1.0000 0.1067
-6.250 -0.3283 0.08044 0.07388 -0.0369 1.0000 0.0629
-6.000 -0.3142 0.07670 0.07016 -0.0379 1.0000 0.0623
-5.750 -0.2980 0.07285 0.06630 -0.0403 1.0000 0.0615
-5.500 -0.2799 0.06890 0.06236 -0.0435 1.0000 0.0603
-5.250 -0.2594 0.06475 0.05820 -0.0474 1.0000 0.0586
-5.000 -0.2355 0.06018 0.05358 -0.0525 1.0000 0.0567
-4.750 -0.2080 0.05516 0.04846 -0.0584 1.0000 0.0551
-4.500 -0.1771 0.04970 0.04280 -0.0649 1.0000 0.0537
-4.250 -0.1443 0.04412 0.03692 -0.0711 1.0000 0.0529
-4.000 -0.1136 0.03936 0.03176 -0.0758 1.0000 0.0530
-3.750 -0.0893 0.03639 0.02848 -0.0780 1.0000 0.0548
-3.500 -0.0382 0.03191 0.02317 -0.0851 0.9910 0.0601
-3.250 0.0109 0.02812 0.01855 -0.0905 0.9824 0.0645
-3.000 0.0546 0.02636 0.01642 -0.0943 0.9714 0.0747
-2.750 0.0979 0.02473 0.01448 -0.0977 0.9605 0.0861
-2.500 0.1406 0.02340 0.01281 -0.1006 0.9491 0.1047
-2.250 0.1819 0.02261 0.01191 -0.1032 0.9373 0.1327
-2.000 0.2207 0.02200 0.01115 -0.1053 0.9239 0.1720
-1.750 0.2567 0.02161 0.01068 -0.1068 0.9100 0.2123
-1.500 0.2908 0.02135 0.01031 -0.1079 0.8960 0.2533
-1.250 0.3232 0.02103 0.00992 -0.1086 0.8824 0.2863
-1.000 0.3545 0.02070 0.00950 -0.1090 0.8692 0.3157
-0.750 0.3844 0.02037 0.00916 -0.1091 0.8563 0.3449
-0.500 0.4131 0.02011 0.00896 -0.1089 0.8435 0.3818
-0.250 0.4406 0.01969 0.00882 -0.1086 0.8311 0.4515
0.000 0.4603 0.01841 0.00857 -0.1061 0.8176 1.0000
0.250 0.4883 0.01867 0.00848 -0.1057 0.8036 1.0000
0.500 0.5156 0.01893 0.00848 -0.1052 0.7898 1.0000
0.750 0.5425 0.01919 0.00852 -0.1047 0.7761 1.0000
1.000 0.5694 0.01946 0.00862 -0.1041 0.7628 1.0000
1.250 0.5962 0.01975 0.00875 -0.1035 0.7502 1.0000
1.500 0.6231 0.02003 0.00889 -0.1030 0.7387 1.0000
1.750 0.6498 0.02035 0.00915 -0.1025 0.7264 1.0000
2.000 0.6766 0.02071 0.00945 -0.1022 0.7140 1.0000
2.250 0.7033 0.02106 0.00977 -0.1018 0.7023 1.0000
2.500 0.7301 0.02140 0.01008 -0.1014 0.6914 1.0000
2.750 0.7569 0.02176 0.01046 -0.1010 0.6798 1.0000
3.000 0.7836 0.02217 0.01091 -0.1007 0.6674 1.0000
3.250 0.8101 0.02257 0.01135 -0.1004 0.6555 1.0000
3.500 0.8367 0.02293 0.01175 -0.0999 0.6444 1.0000
3.750 0.8632 0.02328 0.01221 -0.0994 0.6337 1.0000
4.000 0.8895 0.02376 0.01280 -0.0992 0.6218 1.0000
4.250 0.9157 0.02420 0.01336 -0.0988 0.6108 1.0000
4.500 0.9420 0.02455 0.01380 -0.0982 0.6004 1.0000
4.750 0.9679 0.02496 0.01441 -0.0977 0.5886 1.0000
5.000 0.9934 0.02546 0.01510 -0.0973 0.5765 1.0000
5.250 1.0189 0.02588 0.01570 -0.0966 0.5643 1.0000
5.500 1.0429 0.02576 0.01566 -0.0949 0.5430 1.0000
5.750 1.0648 0.02540 0.01536 -0.0927 0.5105 1.0000
6.000 1.0858 0.02535 0.01537 -0.0908 0.4719 1.0000
6.250 1.1052 0.02546 0.01535 -0.0886 0.4204 1.0000
6.500 1.1231 0.02609 0.01586 -0.0868 0.3475 1.0000
6.750 1.1300 0.02838 0.01699 -0.0845 0.1828 1.0000
7.000 1.1373 0.03139 0.01945 -0.0829 0.1251 1.0000
7.250 1.1491 0.03356 0.02158 -0.0815 0.1039 1.0000
7.500 1.1593 0.03567 0.02365 -0.0799 0.0936 1.0000
7.750 1.1705 0.03753 0.02566 -0.0782 0.0874 1.0000
8.000 1.1796 0.03952 0.02772 -0.0763 0.0833 1.0000
8.250 1.1897 0.04144 0.02973 -0.0743 0.0802 1.0000
8.500 1.2044 0.04308 0.03159 -0.0724 0.0773 1.0000
8.750 1.2224 0.04470 0.03342 -0.0707 0.0741 1.0000
9.000 1.2419 0.04644 0.03520 -0.0694 0.0702 1.0000
9.250 1.2666 0.04843 0.03724 -0.0684 0.0664 1.0000
9.500 1.2928 0.05053 0.03967 -0.0675 0.0640 1.0000
9.750 1.3173 0.05311 0.04258 -0.0667 0.0621 1.0000
10.000 1.3377 0.05607 0.04593 -0.0658 0.0604 1.0000
10.250 1.3522 0.05912 0.04927 -0.0646 0.0582 1.0000
10.500 1.3659 0.06251 0.05274 -0.0639 0.0553 1.0000
10.750 1.3674 0.06622 0.05687 -0.0620 0.0537 1.0000
11.000 1.3630 0.06996 0.06110 -0.0597 0.0529 1.0000
11.250 1.3515 0.07362 0.06518 -0.0573 0.0520 1.0000
11.500 1.3352 0.07739 0.06929 -0.0550 0.0514 1.0000
11.750 1.3172 0.08170 0.07391 -0.0540 0.0508 1.0000
12.000 1.2975 0.08661 0.07911 -0.0541 0.0505 1.0000
12.250 1.2762 0.09220 0.08496 -0.0554 0.0504 1.0000
12.500 1.2535 0.09852 0.09152 -0.0577 0.0506 1.0000
12.750 1.2297 0.10566 0.09886 -0.0610 0.0510 1.0000
13.000 1.2056 0.11359 0.10696 -0.0654 0.0515 1.0000
|
Polar data table (+)
Polar graphs
<< Back to ONERA HOR07 AIRFOIL (hor07-il)