NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 1,000,000 Max Cl/Cd: 91.12 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08n1-il-1000000.txt Download as CSV file: xf-rc08n1-il-1000000.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 = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4890 0.08101 0.07970 0.0061 1.0000 0.0077
-8.000 -0.4948 0.07683 0.07537 0.0048 0.8992 0.0077
-7.750 -0.5021 0.07229 0.07073 0.0022 0.8694 0.0077
-7.500 -0.5128 0.06720 0.06556 -0.0026 0.8516 0.0077
-7.250 -0.5160 0.06215 0.06042 -0.0052 0.8402 0.0077
-7.000 -0.5145 0.05716 0.05532 -0.0071 0.8310 0.0077
-6.750 -0.5098 0.05217 0.05022 -0.0086 0.8234 0.0077
-6.500 -0.5028 0.04720 0.04511 -0.0095 0.8167 0.0077
-6.250 -0.4931 0.04235 0.04013 -0.0100 0.8103 0.0077
-6.000 -0.4817 0.03765 0.03526 -0.0101 0.8046 0.0078
-5.750 -0.4684 0.03308 0.03052 -0.0099 0.7992 0.0078
-5.500 -0.4538 0.02863 0.02588 -0.0094 0.7939 0.0078
-5.250 -0.4380 0.02434 0.02135 -0.0086 0.7892 0.0078
-5.000 -0.4536 0.03179 0.02823 -0.0052 0.7954 0.0082
-4.750 -0.4332 0.03027 0.02657 -0.0053 0.7899 0.0086
-4.500 -0.4105 0.02857 0.02473 -0.0051 0.7847 0.0094
-4.250 -0.3821 0.02720 0.02316 -0.0040 0.7794 0.0115
-4.000 -0.3558 0.02561 0.02132 -0.0031 0.7746 0.0119
-3.750 -0.3309 0.02351 0.01895 -0.0022 0.7698 0.0120
-3.500 -0.3063 0.02140 0.01656 -0.0013 0.7648 0.0120
-3.250 -0.2813 0.01941 0.01428 -0.0005 0.7602 0.0120
-3.000 -0.2587 0.01496 0.00952 0.0005 0.7556 0.0134
-2.750 -0.2323 0.01424 0.00871 0.0006 0.7506 0.0149
-2.500 -0.2041 0.01452 0.00893 0.0006 0.7458 0.0179
-2.250 -0.1762 0.01403 0.00831 0.0009 0.7406 0.0182
-2.000 -0.1509 0.01141 0.00553 0.0015 0.7359 0.0209
-1.750 -0.1240 0.01074 0.00482 0.0017 0.7310 0.0228
-1.500 -0.0971 0.00999 0.00404 0.0021 0.7253 0.0239
-1.250 -0.0711 0.00910 0.00300 0.0030 0.7199 0.0202
-1.000 -0.0440 0.00870 0.00257 0.0033 0.7139 0.0198
-0.750 -0.0169 0.00838 0.00220 0.0035 0.7078 0.0196
-0.500 0.0104 0.00808 0.00185 0.0036 0.7016 0.0199
-0.250 0.0381 0.00788 0.00162 0.0037 0.6948 0.0192
0.000 0.0659 0.00770 0.00139 0.0038 0.6883 0.0190
0.250 0.0937 0.00753 0.00117 0.0038 0.6813 0.0196
0.500 0.1217 0.00743 0.00100 0.0039 0.6750 0.0215
0.750 0.1498 0.00736 0.00091 0.0038 0.6677 0.0234
1.000 0.1778 0.00731 0.00086 0.0038 0.6606 0.0263
1.250 0.2058 0.00723 0.00083 0.0037 0.6527 0.0521
1.500 0.2266 0.00580 0.00076 0.0044 0.6433 0.5597
1.750 0.2418 0.00460 0.00083 0.0074 0.6328 0.9386
2.000 0.3034 0.00477 0.00101 0.0000 0.6146 0.9905
2.250 0.3555 0.00491 0.00109 -0.0055 0.5971 0.9991
2.500 0.3865 0.00498 0.00111 -0.0064 0.5814 1.0000
2.750 0.4137 0.00505 0.00112 -0.0063 0.5616 1.0000
3.000 0.4408 0.00514 0.00115 -0.0063 0.5415 1.0000
3.250 0.4679 0.00525 0.00120 -0.0063 0.5203 1.0000
3.500 0.4948 0.00543 0.00128 -0.0063 0.4852 1.0000
4.000 0.5477 0.00611 0.00154 -0.0064 0.3714 1.0000
4.250 0.5737 0.00653 0.00174 -0.0063 0.3120 1.0000
4.500 0.5990 0.00705 0.00199 -0.0063 0.2437 1.0000
4.750 0.6239 0.00759 0.00227 -0.0062 0.1803 1.0000
5.000 0.6487 0.00808 0.00255 -0.0060 0.1322 1.0000
5.250 0.6736 0.00846 0.00281 -0.0057 0.1035 1.0000
5.500 0.6985 0.00881 0.00306 -0.0054 0.0815 1.0000
5.750 0.7234 0.00915 0.00335 -0.0051 0.0648 1.0000
6.000 0.7481 0.00951 0.00366 -0.0047 0.0500 1.0000
6.250 0.7726 0.00990 0.00399 -0.0044 0.0373 1.0000
6.500 0.7972 0.01025 0.00432 -0.0040 0.0299 1.0000
6.750 0.8218 0.01062 0.00473 -0.0036 0.0239 1.0000
7.000 0.8453 0.01119 0.00530 -0.0030 0.0178 1.0000
7.250 0.8699 0.01152 0.00567 -0.0026 0.0159 1.0000
7.500 0.8940 0.01189 0.00604 -0.0022 0.0133 1.0000
7.750 0.9162 0.01267 0.00690 -0.0015 0.0106 1.0000
8.000 0.9403 0.01303 0.00730 -0.0011 0.0094 1.0000
8.250 0.9637 0.01352 0.00784 -0.0006 0.0081 1.0000
8.500 0.9823 0.01488 0.00936 0.0006 0.0066 1.0000
8.750 1.0056 0.01538 0.00993 0.0011 0.0061 1.0000
9.000 1.0275 0.01612 0.01076 0.0017 0.0055 1.0000
9.250 1.0490 0.01689 0.01161 0.0023 0.0050 1.0000
9.500 1.0704 0.01764 0.01244 0.0029 0.0046 1.0000
9.750 1.0873 0.01913 0.01408 0.0040 0.0042 1.0000
10.000 1.0951 0.02220 0.01749 0.0060 0.0039 1.0000
10.250 1.1128 0.02344 0.01890 0.0069 0.0039 1.0000
10.500 1.1286 0.02492 0.02056 0.0079 0.0038 1.0000
10.750 1.1421 0.02663 0.02247 0.0090 0.0037 1.0000
11.000 1.1530 0.02855 0.02460 0.0102 0.0037 1.0000
11.250 1.1604 0.03069 0.02697 0.0116 0.0036 1.0000
11.500 1.1626 0.03298 0.02948 0.0134 0.0035 1.0000
11.750 1.1580 0.03555 0.03225 0.0151 0.0035 1.0000
12.000 1.1511 0.03867 0.03559 0.0157 0.0034 1.0000
12.250 1.1414 0.04242 0.03955 0.0155 0.0034 1.0000
12.500 1.1299 0.04672 0.04405 0.0145 0.0034 1.0000
12.750 1.1162 0.05169 0.04921 0.0126 0.0034 1.0000
13.000 1.0999 0.05757 0.05527 0.0096 0.0034 1.0000
13.250 1.0787 0.06499 0.06288 0.0051 0.0034 1.0000
13.500 1.0569 0.07347 0.07152 -0.0004 0.0034 1.0000
13.750 1.0314 0.08400 0.08219 -0.0074 0.0035 1.0000
14.000 0.9992 0.09731 0.09564 -0.0153 0.0036 1.0000
14.250 0.9232 0.12219 0.12063 -0.0270 0.0039 1.0000
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Polar data table (+)
Polar graphs
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