NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 50,000 Max Cl/Cd: 33.71 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc08n1-il-50000-n5.txt Download as CSV file: xf-rc08n1-il-50000-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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5885 0.10108 0.09499 0.0082 1.0000 0.1137
-8.000 -0.5841 0.09721 0.09116 0.0068 1.0000 0.1198
-7.750 -0.5998 0.09388 0.08783 -0.0031 1.0000 0.1242
-7.500 -0.5809 0.08888 0.08293 0.0014 1.0000 0.1309
-6.750 -0.5712 0.07709 0.07112 -0.0062 1.0000 0.1644
-5.750 -0.4995 0.05706 0.04994 -0.0167 1.0000 0.0804
-5.500 -0.4743 0.05259 0.04482 -0.0167 1.0000 0.0554
-5.250 -0.4570 0.04846 0.04061 -0.0167 1.0000 0.0522
-5.000 -0.4374 0.04501 0.03683 -0.0164 1.0000 0.0499
-4.750 -0.4169 0.04208 0.03345 -0.0160 1.0000 0.0506
-4.500 -0.3957 0.03939 0.03037 -0.0154 1.0000 0.0516
-4.250 -0.3740 0.03677 0.02737 -0.0146 1.0000 0.0508
-4.000 -0.3523 0.03435 0.02455 -0.0135 1.0000 0.0493
-3.750 -0.3313 0.03223 0.02207 -0.0122 1.0000 0.0483
-3.500 -0.3115 0.03040 0.01990 -0.0107 1.0000 0.0477
-3.250 -0.2922 0.02881 0.01801 -0.0091 1.0000 0.0475
-3.000 -0.2725 0.02739 0.01633 -0.0076 1.0000 0.0476
-2.750 -0.2522 0.02613 0.01488 -0.0063 1.0000 0.0485
-2.500 -0.2152 0.02484 0.01329 -0.0079 0.9912 0.0530
-2.250 -0.1757 0.02369 0.01186 -0.0098 0.9829 0.0556
-2.000 -0.1374 0.02249 0.01048 -0.0118 0.9745 0.0576
-1.750 -0.1018 0.02155 0.00937 -0.0136 0.9653 0.0611
-1.500 -0.0647 0.02086 0.00843 -0.0157 0.9565 0.0666
-1.250 -0.0295 0.02022 0.00762 -0.0174 0.9468 0.0780
-1.000 0.0547 0.01648 0.00697 -0.0269 0.9550 1.0000
-0.750 0.0917 0.01658 0.00668 -0.0291 0.9436 1.0000
-0.500 0.1257 0.01672 0.00650 -0.0308 0.9318 1.0000
-0.250 0.1576 0.01689 0.00645 -0.0320 0.9199 1.0000
0.000 0.1878 0.01709 0.00647 -0.0328 0.9083 1.0000
0.250 0.2168 0.01731 0.00655 -0.0333 0.8969 1.0000
0.500 0.2448 0.01755 0.00667 -0.0336 0.8858 1.0000
0.750 0.2720 0.01781 0.00683 -0.0336 0.8751 1.0000
1.000 0.2967 0.01809 0.00706 -0.0331 0.8632 1.0000
1.250 0.3210 0.01839 0.00732 -0.0326 0.8513 1.0000
1.500 0.3449 0.01869 0.00760 -0.0319 0.8395 1.0000
1.750 0.3686 0.01900 0.00793 -0.0311 0.8276 1.0000
2.000 0.3920 0.01931 0.00825 -0.0302 0.8157 1.0000
2.250 0.4153 0.01961 0.00859 -0.0292 0.8036 1.0000
2.500 0.4383 0.01991 0.00893 -0.0280 0.7912 1.0000
2.750 0.4611 0.02019 0.00934 -0.0268 0.7784 1.0000
3.000 0.4837 0.02048 0.00972 -0.0256 0.7648 1.0000
3.250 0.5065 0.02075 0.01009 -0.0244 0.7508 1.0000
3.500 0.5288 0.02102 0.01050 -0.0231 0.7353 1.0000
3.750 0.5509 0.02124 0.01087 -0.0217 0.7183 1.0000
4.000 0.5725 0.02128 0.01112 -0.0197 0.6987 1.0000
4.250 0.5932 0.02118 0.01116 -0.0173 0.6739 1.0000
4.500 0.6138 0.02097 0.01108 -0.0147 0.6459 1.0000
4.750 0.6348 0.02077 0.01102 -0.0123 0.6151 1.0000
5.000 0.6556 0.02064 0.01104 -0.0100 0.5776 1.0000
5.250 0.6762 0.02053 0.01107 -0.0075 0.5296 1.0000
5.500 0.6957 0.02064 0.01111 -0.0051 0.4589 1.0000
5.750 0.7124 0.02130 0.01131 -0.0026 0.3582 1.0000
6.000 0.7266 0.02284 0.01222 -0.0009 0.2561 1.0000
6.250 0.7411 0.02465 0.01358 0.0002 0.1894 1.0000
6.500 0.7566 0.02639 0.01506 0.0013 0.1489 1.0000
6.750 0.7733 0.02809 0.01670 0.0026 0.1241 1.0000
7.000 0.7903 0.02971 0.01834 0.0037 0.1022 1.0000
7.250 0.8098 0.03148 0.02019 0.0051 0.0892 1.0000
7.500 0.8297 0.03324 0.02203 0.0062 0.0767 1.0000
7.750 0.8494 0.03499 0.02395 0.0072 0.0658 1.0000
8.000 0.8730 0.03758 0.02692 0.0084 0.0600 1.0000
8.250 0.8928 0.04006 0.02966 0.0094 0.0545 1.0000
8.500 0.9087 0.04283 0.03287 0.0103 0.0487 1.0000
8.750 0.9232 0.04605 0.03657 0.0114 0.0450 1.0000
9.000 0.9346 0.04956 0.04048 0.0125 0.0429 1.0000
9.250 0.9426 0.05321 0.04446 0.0134 0.0414 1.0000
9.500 0.9465 0.05740 0.04883 0.0141 0.0400 1.0000
9.750 0.9409 0.06163 0.05360 0.0151 0.0393 1.0000
10.000 0.9302 0.06598 0.05834 0.0156 0.0387 1.0000
10.250 0.9151 0.07021 0.06283 0.0158 0.0385 1.0000
10.500 0.8981 0.07478 0.06757 0.0148 0.0386 1.0000
10.750 0.8811 0.07997 0.07290 0.0126 0.0389 1.0000
11.000 0.8647 0.08578 0.07879 0.0095 0.0392 1.0000
11.250 0.8495 0.09213 0.08520 0.0057 0.0395 1.0000
11.500 0.8361 0.09891 0.09200 0.0015 0.0398 1.0000
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Polar data table (+)
Polar graphs
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