NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC10-64C AIRFOIL (rc1064c-il) Reynolds number: 100,000 Max Cl/Cd: 45.4 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc1064c-il-100000.txt Download as CSV file: xf-rc1064c-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC10-64C AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.750 -0.5552 0.11847 0.11313 -0.0104 1.0000 0.1086
-10.500 -0.5689 0.11467 0.10940 -0.0152 1.0000 0.1130
-10.250 -0.6031 0.11016 0.10503 -0.0247 1.0000 0.1142
-10.000 -0.5671 0.10569 0.10049 -0.0172 1.0000 0.1171
-9.750 -0.5565 0.10263 0.09743 -0.0162 1.0000 0.1218
-9.500 -0.5684 0.09837 0.09324 -0.0202 1.0000 0.1270
-9.250 -0.6054 0.09275 0.08772 -0.0289 1.0000 0.1286
-9.000 -0.5711 0.08991 0.08486 -0.0223 1.0000 0.1343
-8.750 -0.5770 0.08589 0.08089 -0.0243 1.0000 0.1393
-8.500 -0.6099 0.08145 0.07649 -0.0275 1.0000 0.1423
-8.250 -0.6586 0.07874 0.07352 -0.0284 1.0000 0.1441
-8.000 -0.6112 0.07405 0.06910 -0.0258 1.0000 0.1508
-7.750 -0.6608 0.07219 0.06685 -0.0252 1.0000 0.1581
-7.500 -0.6272 0.06704 0.06201 -0.0239 1.0000 0.1622
-7.250 -0.6456 0.06439 0.05912 -0.0221 1.0000 0.1720
-7.000 -0.6304 0.06081 0.05566 -0.0203 1.0000 0.1759
-6.750 -0.6581 0.04917 0.04229 -0.0168 1.0000 0.0998
-6.500 -0.6469 0.04217 0.03499 -0.0145 1.0000 0.0825
-6.250 -0.6368 0.03856 0.03109 -0.0120 1.0000 0.0796
-6.000 -0.6249 0.03538 0.02748 -0.0093 1.0000 0.0784
-5.750 -0.6103 0.03292 0.02463 -0.0070 1.0000 0.0796
-5.500 -0.5934 0.03051 0.02180 -0.0049 1.0000 0.0801
-5.250 -0.5742 0.02826 0.01916 -0.0030 1.0000 0.0806
-5.000 -0.5532 0.02640 0.01693 -0.0014 1.0000 0.0822
-4.750 -0.5315 0.02505 0.01522 0.0001 1.0000 0.0851
-4.500 -0.5095 0.02340 0.01363 0.0009 1.0000 0.0912
-4.250 -0.4864 0.02226 0.01232 0.0021 1.0000 0.0974
-4.000 -0.4634 0.02090 0.01102 0.0030 1.0000 0.1055
-3.750 -0.4416 0.01992 0.01010 0.0042 1.0000 0.1197
-3.500 -0.4201 0.01906 0.00933 0.0053 1.0000 0.1399
-3.250 -0.3991 0.01870 0.00904 0.0065 1.0000 0.1691
-3.000 -0.3792 0.01733 0.00796 0.0077 1.0000 0.2143
-2.750 -0.3612 0.01572 0.00730 0.0090 1.0000 0.3330
-2.500 -0.3353 0.01456 0.00845 0.0136 1.0000 0.8835
-2.250 -0.2057 0.01608 0.00945 -0.0028 1.0000 0.9701
-2.000 -0.0601 0.01612 0.00900 -0.0253 1.0000 1.0000
-1.750 -0.0571 0.01604 0.00885 -0.0218 1.0000 1.0000
-1.500 -0.0558 0.01600 0.00876 -0.0179 1.0000 1.0000
-1.250 -0.0554 0.01600 0.00871 -0.0139 1.0000 1.0000
-1.000 -0.0551 0.01603 0.00870 -0.0099 1.0000 1.0000
-0.750 -0.0542 0.01609 0.00872 -0.0060 1.0000 1.0000
-0.500 -0.0521 0.01619 0.00877 -0.0023 1.0000 1.0000
-0.250 -0.0485 0.01633 0.00886 0.0011 1.0000 1.0000
0.000 -0.0429 0.01651 0.00899 0.0041 1.0000 1.0000
0.250 -0.0319 0.01675 0.00917 0.0060 0.9993 1.0000
0.500 0.0166 0.01713 0.00952 0.0009 0.9911 1.0000
0.750 0.0634 0.01747 0.00984 -0.0039 0.9823 1.0000
1.000 0.1098 0.01777 0.01013 -0.0084 0.9733 1.0000
1.250 0.1593 0.01801 0.01041 -0.0134 0.9636 1.0000
1.500 0.2135 0.01819 0.01063 -0.0191 0.9532 1.0000
1.750 0.2715 0.01831 0.01084 -0.0254 0.9441 1.0000
2.000 0.3309 0.01837 0.01103 -0.0317 0.9342 1.0000
2.250 0.3981 0.01813 0.01095 -0.0390 0.9196 1.0000
2.500 0.4406 0.01779 0.01075 -0.0410 0.8971 1.0000
2.750 0.4810 0.01713 0.01023 -0.0423 0.8754 1.0000
3.000 0.5236 0.01624 0.00949 -0.0433 0.8552 1.0000
3.250 0.5631 0.01529 0.00864 -0.0434 0.8326 1.0000
3.500 0.5842 0.01489 0.00830 -0.0405 0.7977 1.0000
3.750 0.6077 0.01443 0.00783 -0.0379 0.7565 1.0000
4.000 0.6275 0.01425 0.00761 -0.0350 0.7063 1.0000
4.250 0.6461 0.01423 0.00745 -0.0319 0.6390 1.0000
4.500 0.6602 0.01461 0.00735 -0.0281 0.5236 1.0000
4.750 0.6607 0.01628 0.00769 -0.0227 0.3043 1.0000
5.000 0.6650 0.01826 0.00873 -0.0187 0.1897 1.0000
5.250 0.6775 0.01960 0.00975 -0.0159 0.1547 1.0000
5.500 0.6935 0.02073 0.01070 -0.0137 0.1363 1.0000
5.750 0.7122 0.02182 0.01171 -0.0120 0.1232 1.0000
6.000 0.7339 0.02318 0.01298 -0.0108 0.1151 1.0000
6.250 0.7561 0.02430 0.01415 -0.0096 0.1077 1.0000
6.500 0.7793 0.02581 0.01564 -0.0088 0.1014 1.0000
6.750 0.8026 0.02718 0.01719 -0.0077 0.0968 1.0000
7.000 0.8255 0.02875 0.01886 -0.0067 0.0932 1.0000
7.250 0.8485 0.03128 0.02140 -0.0062 0.0895 1.0000
7.500 0.8657 0.03267 0.02324 -0.0040 0.0866 1.0000
7.750 0.8830 0.03487 0.02583 -0.0020 0.0851 1.0000
8.000 0.8982 0.03743 0.02880 0.0002 0.0842 1.0000
8.250 0.9106 0.04005 0.03181 0.0026 0.0828 1.0000
8.500 0.9220 0.04251 0.03458 0.0048 0.0808 1.0000
8.750 0.9293 0.04568 0.03814 0.0075 0.0804 1.0000
9.000 0.9274 0.05020 0.04324 0.0110 0.0826 1.0000
9.250 0.9240 0.05488 0.04832 0.0140 0.0853 1.0000
9.500 0.9234 0.05953 0.05320 0.0162 0.0876 1.0000
9.750 0.9325 0.06539 0.05908 0.0167 0.0894 1.0000
10.000 0.7775 0.06238 0.05719 0.0301 0.1069 1.0000
10.250 0.7053 0.07011 0.06509 0.0297 0.1096 1.0000
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Polar data table (+)
Polar graphs
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