NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 200,000 Max Cl/Cd: 63.62 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08n1-il-200000.txt Download as CSV file: xf-rc08n1-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5830 0.09523 0.09210 0.0142 1.0000 0.0221
-8.000 -0.5813 0.09092 0.08783 0.0109 1.0000 0.0225
-7.750 -0.5804 0.08632 0.08326 0.0063 1.0000 0.0228
-7.500 -0.5770 0.08148 0.07842 0.0015 1.0000 0.0232
-7.250 -0.5703 0.07660 0.07350 -0.0028 1.0000 0.0238
-7.000 -0.5612 0.07182 0.06865 -0.0066 1.0000 0.0245
-6.750 -0.5475 0.06722 0.06389 -0.0102 1.0000 0.0257
-6.500 -0.5272 0.06429 0.06057 -0.0129 1.0000 0.0268
-6.250 -0.5208 0.05759 0.05380 -0.0142 1.0000 0.0276
-6.000 -0.5096 0.05375 0.05004 -0.0146 1.0000 0.0292
-5.750 -0.4927 0.05068 0.04688 -0.0153 1.0000 0.0324
-5.500 -0.4678 0.04819 0.04372 -0.0156 1.0000 0.0379
-5.250 -0.4533 0.04346 0.03918 -0.0167 1.0000 0.0405
-5.000 -0.4326 0.04077 0.03636 -0.0172 1.0000 0.0447
-4.750 -0.4094 0.03952 0.03441 -0.0164 0.9963 0.0503
-4.500 -0.3804 0.03453 0.02957 -0.0196 0.9819 0.0542
-4.250 -0.3455 0.03235 0.02680 -0.0212 0.9681 0.0643
-4.000 -0.3164 0.02943 0.02387 -0.0228 0.9557 0.0721
-3.750 -0.2892 0.02798 0.02204 -0.0228 0.9427 0.0882
-3.500 -0.2673 0.02552 0.01934 -0.0225 0.9304 0.1064
-2.500 -0.1532 0.01843 0.01077 -0.0161 0.8922 0.0480
-2.250 -0.1276 0.01668 0.00887 -0.0149 0.8845 0.0440
-2.000 -0.1014 0.01561 0.00767 -0.0140 0.8759 0.0434
-1.750 -0.0762 0.01491 0.00690 -0.0131 0.8681 0.0454
-1.500 -0.0518 0.01410 0.00605 -0.0120 0.8602 0.0445
-1.250 -0.0272 0.01348 0.00540 -0.0112 0.8524 0.0445
-1.000 -0.0031 0.01302 0.00487 -0.0102 0.8452 0.0455
-0.750 0.0228 0.01270 0.00446 -0.0097 0.8372 0.0479
-0.500 0.0476 0.01243 0.00407 -0.0088 0.8305 0.0545
-0.250 0.1029 0.00944 0.00410 -0.0137 0.8250 0.9793
0.000 0.1546 0.00950 0.00393 -0.0187 0.8182 1.0000
0.250 0.1804 0.00957 0.00389 -0.0184 0.8095 1.0000
0.500 0.2049 0.00966 0.00384 -0.0176 0.8020 1.0000
0.750 0.2308 0.00972 0.00385 -0.0173 0.7926 1.0000
1.000 0.2561 0.00981 0.00387 -0.0168 0.7841 1.0000
1.250 0.2809 0.00988 0.00386 -0.0161 0.7759 1.0000
1.500 0.3068 0.00995 0.00391 -0.0157 0.7664 1.0000
1.750 0.3319 0.01003 0.00394 -0.0151 0.7582 1.0000
2.000 0.3572 0.01009 0.00398 -0.0146 0.7492 1.0000
2.250 0.3829 0.01016 0.00408 -0.0141 0.7396 1.0000
2.500 0.4079 0.01021 0.00410 -0.0134 0.7300 1.0000
2.750 0.4324 0.01017 0.00404 -0.0125 0.7171 1.0000
3.000 0.4568 0.01009 0.00396 -0.0115 0.7014 1.0000
3.250 0.4817 0.01005 0.00395 -0.0106 0.6866 1.0000
3.500 0.5069 0.01001 0.00393 -0.0099 0.6715 1.0000
3.750 0.5320 0.00998 0.00392 -0.0091 0.6549 1.0000
4.000 0.5570 0.00996 0.00389 -0.0083 0.6365 1.0000
4.250 0.5823 0.00996 0.00395 -0.0076 0.6134 1.0000
4.500 0.6073 0.01000 0.00403 -0.0069 0.5867 1.0000
4.750 0.6319 0.01009 0.00409 -0.0061 0.5504 1.0000
5.000 0.6559 0.01031 0.00419 -0.0052 0.4958 1.0000
5.250 0.6783 0.01084 0.00439 -0.0043 0.4074 1.0000
5.500 0.6984 0.01194 0.00492 -0.0036 0.2849 1.0000
5.750 0.7173 0.01340 0.00576 -0.0030 0.1701 1.0000
6.000 0.7378 0.01459 0.00667 -0.0024 0.1215 1.0000
6.250 0.7590 0.01561 0.00758 -0.0017 0.0918 1.0000
6.500 0.7786 0.01690 0.00875 -0.0008 0.0730 1.0000
6.750 0.7995 0.01798 0.00983 0.0001 0.0595 1.0000
7.000 0.8209 0.01900 0.01095 0.0010 0.0505 1.0000
7.250 0.8410 0.02048 0.01252 0.0022 0.0438 1.0000
7.500 0.8614 0.02184 0.01388 0.0030 0.0377 1.0000
7.750 0.8821 0.02399 0.01624 0.0043 0.0344 1.0000
8.000 0.9037 0.02506 0.01747 0.0051 0.0297 1.0000
8.250 0.9230 0.02738 0.01990 0.0060 0.0270 1.0000
8.500 0.9380 0.03211 0.02510 0.0074 0.0256 1.0000
8.750 0.9545 0.03461 0.02809 0.0088 0.0248 1.0000
9.000 0.9689 0.03702 0.03095 0.0102 0.0230 1.0000
9.250 0.9782 0.04039 0.03476 0.0117 0.0218 1.0000
9.500 0.9794 0.04515 0.04002 0.0134 0.0220 1.0000
9.750 0.9740 0.05018 0.04548 0.0150 0.0225 1.0000
10.000 0.9634 0.05506 0.05070 0.0162 0.0230 1.0000
10.250 0.9474 0.05937 0.05524 0.0173 0.0235 1.0000
10.500 0.9289 0.06375 0.05980 0.0169 0.0238 1.0000
10.750 0.9097 0.06880 0.06500 0.0150 0.0241 1.0000
11.000 0.8903 0.07462 0.07095 0.0117 0.0244 1.0000
11.250 0.8704 0.08150 0.07794 0.0069 0.0246 1.0000
11.500 0.8511 0.08952 0.08603 0.0012 0.0250 1.0000
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Polar data table (+)
Polar graphs
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