NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 500,000 Max Cl/Cd: 67.6 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08n1-il-500000-n5.txt Download as CSV file: xf-rc08n1-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-08(N)1 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.6012 0.09264 0.09065 0.0181 1.0000 0.0059
-8.000 -0.6015 0.08792 0.08597 0.0143 1.0000 0.0055
-7.750 -0.6028 0.08267 0.08074 0.0088 1.0000 0.0053
-7.500 -0.6009 0.07724 0.07514 0.0039 0.8910 0.0050
-7.250 -0.5977 0.07177 0.06950 0.0004 0.8634 0.0048
-6.750 -0.5809 0.06121 0.05858 -0.0047 0.8356 0.0047
-6.500 -0.5682 0.05669 0.05389 -0.0060 0.8260 0.0048
-6.250 -0.5535 0.05241 0.04941 -0.0068 0.8180 0.0050
-6.000 -0.5370 0.04821 0.04500 -0.0072 0.8105 0.0052
-5.750 -0.5191 0.04408 0.04061 -0.0072 0.8040 0.0055
-5.500 -0.4992 0.03994 0.03620 -0.0067 0.7976 0.0061
-5.000 -0.4553 0.03048 0.02596 -0.0039 0.7868 0.0069
-4.750 -0.4353 0.02637 0.02143 -0.0026 0.7815 0.0072
-4.500 -0.4123 0.02487 0.01973 -0.0023 0.7764 0.0075
-4.250 -0.3881 0.02334 0.01797 -0.0019 0.7709 0.0080
-4.000 -0.3631 0.02129 0.01562 -0.0011 0.7660 0.0086
-3.750 -0.3367 0.01912 0.01311 -0.0001 0.7613 0.0098
-3.500 -0.3088 0.01843 0.01205 0.0006 0.7561 0.0107
-3.250 -0.2838 0.01593 0.00928 0.0013 0.7516 0.0112
-3.000 -0.2576 0.01481 0.00809 0.0015 0.7467 0.0119
-2.750 -0.2309 0.01398 0.00717 0.0017 0.7416 0.0125
-2.250 -0.1771 0.01270 0.00570 0.0022 0.7319 0.0142
-2.000 -0.1505 0.01205 0.00496 0.0025 0.7267 0.0145
-1.750 -0.1241 0.01142 0.00426 0.0029 0.7218 0.0145
-1.500 -0.0976 0.01088 0.00367 0.0032 0.7161 0.0145
-1.250 -0.0711 0.01044 0.00317 0.0035 0.7109 0.0148
-1.000 -0.0442 0.01007 0.00275 0.0037 0.7054 0.0152
-0.750 -0.0170 0.00980 0.00244 0.0039 0.6992 0.0158
-0.500 0.0102 0.00954 0.00213 0.0040 0.6928 0.0164
-0.250 0.0374 0.00926 0.00180 0.0041 0.6856 0.0174
0.000 0.0648 0.00907 0.00155 0.0042 0.6787 0.0194
0.250 0.0926 0.00897 0.00141 0.0042 0.6707 0.0197
0.500 0.1204 0.00889 0.00130 0.0042 0.6627 0.0201
0.750 0.1482 0.00884 0.00121 0.0042 0.6546 0.0207
1.000 0.1761 0.00878 0.00115 0.0042 0.6460 0.0220
1.500 0.2317 0.00867 0.00109 0.0041 0.6291 0.0642
1.750 0.2498 0.00697 0.00104 0.0052 0.6205 0.6457
2.000 0.2768 0.00611 0.00118 0.0059 0.6080 0.9440
2.250 0.3270 0.00626 0.00128 0.0010 0.5836 0.9795
2.500 0.3787 0.00645 0.00136 -0.0044 0.5528 0.9982
2.750 0.4103 0.00658 0.00141 -0.0054 0.5269 1.0000
3.000 0.4369 0.00674 0.00148 -0.0053 0.4951 1.0000
3.250 0.4632 0.00697 0.00156 -0.0052 0.4537 1.0000
3.500 0.4894 0.00724 0.00168 -0.0051 0.4112 1.0000
3.750 0.5152 0.00764 0.00186 -0.0051 0.3556 1.0000
4.000 0.5405 0.00812 0.00211 -0.0050 0.2929 1.0000
4.250 0.5656 0.00862 0.00237 -0.0049 0.2331 1.0000
4.500 0.5905 0.00908 0.00264 -0.0047 0.1854 1.0000
4.750 0.6152 0.00956 0.00294 -0.0045 0.1417 1.0000
5.000 0.6399 0.00998 0.00323 -0.0043 0.1124 1.0000
5.250 0.6646 0.01037 0.00356 -0.0040 0.0890 1.0000
5.500 0.6893 0.01074 0.00387 -0.0036 0.0714 1.0000
5.750 0.7141 0.01108 0.00420 -0.0033 0.0590 1.0000
6.000 0.7386 0.01146 0.00455 -0.0029 0.0475 1.0000
6.250 0.7629 0.01187 0.00493 -0.0026 0.0370 1.0000
6.500 0.7868 0.01232 0.00539 -0.0021 0.0289 1.0000
6.750 0.8111 0.01270 0.00582 -0.0017 0.0247 1.0000
7.000 0.8344 0.01327 0.00640 -0.0012 0.0188 1.0000
7.250 0.8584 0.01367 0.00688 -0.0007 0.0170 1.0000
7.500 0.8820 0.01417 0.00744 -0.0002 0.0148 1.0000
7.750 0.9046 0.01482 0.00815 0.0003 0.0122 1.0000
8.000 0.9270 0.01551 0.00893 0.0009 0.0104 1.0000
8.250 0.9501 0.01603 0.00954 0.0014 0.0091 1.0000
8.500 0.9732 0.01654 0.01009 0.0018 0.0075 1.0000
8.750 0.9941 0.01742 0.01103 0.0025 0.0059 1.0000
9.000 1.0164 0.01805 0.01176 0.0030 0.0053 1.0000
9.250 1.0376 0.01883 0.01267 0.0036 0.0046 1.0000
9.500 1.0586 0.01962 0.01355 0.0042 0.0040 1.0000
9.750 1.0779 0.02066 0.01470 0.0049 0.0035 1.0000
10.000 1.0938 0.02220 0.01643 0.0060 0.0031 1.0000
10.250 1.1114 0.02340 0.01781 0.0068 0.0029 1.0000
10.500 1.1272 0.02480 0.01941 0.0078 0.0027 1.0000
10.750 1.1410 0.02640 0.02123 0.0089 0.0025 1.0000
11.000 1.1526 0.02816 0.02322 0.0101 0.0024 1.0000
11.250 1.1616 0.03009 0.02538 0.0114 0.0023 1.0000
11.500 1.1664 0.03217 0.02770 0.0129 0.0022 1.0000
11.750 1.1656 0.03447 0.03022 0.0146 0.0021 1.0000
12.000 1.1624 0.03722 0.03320 0.0154 0.0021 1.0000
12.250 1.1570 0.04045 0.03665 0.0156 0.0020 1.0000
12.500 1.1489 0.04424 0.04067 0.0151 0.0020 1.0000
12.750 1.1377 0.04870 0.04536 0.0138 0.0020 1.0000
13.000 1.1236 0.05397 0.05084 0.0117 0.0020 1.0000
13.250 1.1067 0.06018 0.05726 0.0084 0.0020 1.0000
13.500 1.0865 0.06767 0.06495 0.0039 0.0020 1.0000
13.750 1.0637 0.07673 0.07419 -0.0020 0.0021 1.0000
14.000 1.0367 0.08783 0.08546 -0.0089 0.0021 1.0000
14.250 1.0038 0.10123 0.09900 -0.0162 0.0022 1.0000
14.500 0.9636 0.11631 0.11417 -0.0232 0.0023 1.0000
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Polar data table (+)
Polar graphs
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