NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Xfoil prediction polar at RE=500,000 Ncrit=9
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Airfoil: NASA/LANGLEY RC-08(N)1 AIRFOIL (rc08n1-il) Reynolds number: 500,000 Max Cl/Cd: 81.52 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc08n1-il-500000.txt Download as CSV file: xf-rc08n1-il-500000.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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.5938 0.09397 0.09198 0.0178 1.0000 0.0102
-8.000 -0.5924 0.08958 0.08762 0.0144 1.0000 0.0104
-7.750 -0.5921 0.08480 0.08287 0.0094 1.0000 0.0105
-7.500 -0.5878 0.07965 0.07771 0.0039 1.0000 0.0106
-7.250 -0.5798 0.07449 0.07251 -0.0008 1.0000 0.0109
-7.000 -0.5684 0.06948 0.06742 -0.0048 1.0000 0.0112
-6.750 -0.5515 0.06473 0.06255 -0.0082 1.0000 0.0116
-6.500 -0.5338 0.06037 0.05804 -0.0102 1.0000 0.0118
-6.250 -0.5133 0.05602 0.05352 -0.0122 0.9590 0.0119
-6.000 -0.4984 0.05233 0.04956 -0.0119 0.9133 0.0120
-5.750 -0.4837 0.04874 0.04569 -0.0111 0.8918 0.0120
-5.500 -0.4672 0.04507 0.04172 -0.0103 0.8775 0.0120
-5.250 -0.4585 0.03813 0.03445 -0.0098 0.8666 0.0125
-5.000 -0.4420 0.03544 0.03159 -0.0095 0.8572 0.0131
-4.750 -0.4219 0.03310 0.02906 -0.0092 0.8484 0.0138
-4.500 -0.3998 0.03081 0.02653 -0.0085 0.8410 0.0150
-4.250 -0.3707 0.03003 0.02548 -0.0073 0.8337 0.0179
-4.000 -0.3455 0.02848 0.02360 -0.0063 0.8271 0.0182
-3.750 -0.3259 0.02345 0.01809 -0.0051 0.8209 0.0191
-3.500 -0.3027 0.02135 0.01582 -0.0047 0.8151 0.0204
-3.250 -0.2773 0.01991 0.01419 -0.0043 0.8087 0.0223
-3.000 -0.2483 0.02070 0.01469 -0.0035 0.8031 0.0269
-2.750 -0.2253 0.01754 0.01142 -0.0034 0.7973 0.0333
-2.500 -0.1997 0.01673 0.01051 -0.0033 0.7915 0.0440
-2.250 -0.1738 0.01588 0.00956 -0.0032 0.7860 0.0565
-1.750 -0.1150 0.01223 0.00539 -0.0007 0.7756 0.0280
-1.500 -0.0877 0.01167 0.00481 -0.0003 0.7697 0.0267
-1.250 -0.0623 0.01079 0.00388 0.0004 0.7641 0.0263
-1.000 -0.0359 0.01035 0.00339 0.0007 0.7586 0.0271
-0.750 -0.0094 0.00987 0.00284 0.0011 0.7523 0.0267
-0.500 0.0173 0.00954 0.00241 0.0015 0.7467 0.0265
-0.250 0.0448 0.00928 0.00210 0.0016 0.7397 0.0267
0.000 0.0723 0.00913 0.00185 0.0018 0.7338 0.0283
0.250 0.1002 0.00900 0.00172 0.0019 0.7265 0.0341
0.500 0.1273 0.00879 0.00163 0.0020 0.7201 0.0920
0.750 0.1393 0.00650 0.00148 0.0044 0.7132 0.7479
1.000 0.1763 0.00609 0.00162 0.0032 0.7067 0.9523
1.250 0.2338 0.00622 0.00174 -0.0032 0.6989 0.9893
1.500 0.2833 0.00630 0.00175 -0.0080 0.6916 1.0000
1.750 0.3103 0.00629 0.00173 -0.0079 0.6824 1.0000
2.000 0.3369 0.00629 0.00169 -0.0077 0.6710 1.0000
2.250 0.3634 0.00629 0.00167 -0.0074 0.6574 1.0000
2.500 0.3900 0.00631 0.00165 -0.0072 0.6436 1.0000
2.750 0.4167 0.00634 0.00166 -0.0070 0.6301 1.0000
3.000 0.4433 0.00639 0.00168 -0.0068 0.6152 1.0000
3.250 0.4698 0.00645 0.00171 -0.0066 0.5975 1.0000
3.500 0.4963 0.00653 0.00178 -0.0063 0.5782 1.0000
3.750 0.5228 0.00664 0.00185 -0.0061 0.5565 1.0000
4.000 0.5489 0.00680 0.00194 -0.0059 0.5244 1.0000
4.250 0.5747 0.00705 0.00205 -0.0056 0.4768 1.0000
4.500 0.6000 0.00746 0.00225 -0.0054 0.4149 1.0000
4.750 0.6246 0.00804 0.00254 -0.0053 0.3382 1.0000
5.000 0.6485 0.00877 0.00290 -0.0051 0.2534 1.0000
5.250 0.6719 0.00954 0.00332 -0.0049 0.1747 1.0000
5.500 0.6955 0.01021 0.00375 -0.0046 0.1250 1.0000
5.750 0.7197 0.01072 0.00417 -0.0043 0.0953 1.0000
6.000 0.7436 0.01125 0.00460 -0.0039 0.0717 1.0000
6.250 0.7675 0.01176 0.00507 -0.0034 0.0550 1.0000
6.500 0.7911 0.01233 0.00562 -0.0029 0.0427 1.0000
6.750 0.8142 0.01297 0.00627 -0.0023 0.0337 1.0000
7.000 0.8359 0.01386 0.00721 -0.0016 0.0266 1.0000
7.250 0.8594 0.01438 0.00781 -0.0010 0.0232 1.0000
7.500 0.8822 0.01501 0.00848 -0.0004 0.0196 1.0000
7.750 0.9001 0.01656 0.01017 0.0009 0.0167 1.0000
8.000 0.9237 0.01701 0.01070 0.0013 0.0145 1.0000
8.250 0.9461 0.01764 0.01139 0.0019 0.0125 1.0000
8.500 0.9584 0.02040 0.01436 0.0037 0.0108 1.0000
8.750 0.9801 0.02125 0.01536 0.0045 0.0102 1.0000
9.000 0.9993 0.02270 0.01700 0.0055 0.0095 1.0000
9.250 1.0170 0.02448 0.01898 0.0066 0.0088 1.0000
9.500 1.0328 0.02660 0.02136 0.0078 0.0085 1.0000
9.750 1.0456 0.02927 0.02434 0.0092 0.0083 1.0000
10.000 1.0531 0.03277 0.02823 0.0109 0.0082 1.0000
10.250 1.0587 0.03588 0.03168 0.0125 0.0081 1.0000
10.500 1.0651 0.03807 0.03409 0.0137 0.0078 1.0000
10.750 1.0703 0.03985 0.03601 0.0147 0.0074 1.0000
11.000 1.0658 0.04222 0.03853 0.0163 0.0072 1.0000
11.250 1.0536 0.04578 0.04231 0.0171 0.0071 1.0000
11.500 1.0394 0.04995 0.04670 0.0167 0.0071 1.0000
11.750 1.0244 0.05470 0.05166 0.0152 0.0071 1.0000
12.000 1.0083 0.06018 0.05734 0.0126 0.0071 1.0000
12.250 0.8168 0.07004 0.06772 0.0130 0.0093 1.0000
12.500 0.7928 0.07882 0.07660 0.0087 0.0094 1.0000
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Polar data table (+)
Polar graphs
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