NASA RC(5)-10 AIRFOIL (rc510-il) Xfoil prediction polar at RE=50,000 Ncrit=9
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Airfoil: NASA RC(5)-10 AIRFOIL (rc510-il) Reynolds number: 50,000 Max Cl/Cd: 31.62 at α=5.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rc510-il-50000.txt Download as CSV file: xf-rc510-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA RC(5)-10 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5382 0.12215 0.11570 0.0077 1.0000 0.1830
-9.250 -0.5080 0.11590 0.10942 0.0109 1.0000 0.1944
-9.000 -0.5324 0.11473 0.10839 0.0052 1.0000 0.1986
-8.750 -0.5065 0.10926 0.10291 0.0080 1.0000 0.2120
-8.500 -0.4960 0.10504 0.09872 0.0082 1.0000 0.2211
-8.250 -0.5199 0.10337 0.09722 0.0029 1.0000 0.2297
-8.000 -0.5016 0.09889 0.09274 0.0047 1.0000 0.2441
-7.750 -0.4900 0.09493 0.08883 0.0053 1.0000 0.2579
-7.500 -0.4868 0.09152 0.08550 0.0050 1.0000 0.2736
-7.250 -0.4936 0.08862 0.08271 0.0036 1.0000 0.2913
-7.000 -0.4727 0.08472 0.07884 0.0064 1.0000 0.3148
-6.250 -0.4256 0.07665 0.07090 0.0179 1.0000 0.4533
-6.000 -0.3957 0.07454 0.06879 0.0246 1.0000 0.5373
-5.750 -0.3388 0.07057 0.06473 0.0295 1.0000 0.6412
-5.500 -0.2618 0.06499 0.05903 0.0299 1.0000 0.7823
-4.750 -0.3118 0.05865 0.05312 0.0272 1.0000 0.6774
-4.500 -0.3852 0.05774 0.05256 0.0288 1.0000 0.6052
-4.250 -0.4386 0.05762 0.05263 0.0355 1.0000 0.6017
-4.000 -0.4273 0.04398 0.03629 -0.0106 1.0000 0.1956
-3.750 -0.4035 0.04098 0.03260 -0.0101 1.0000 0.1643
-3.500 -0.3815 0.03876 0.02970 -0.0091 1.0000 0.1512
-3.250 -0.3613 0.03663 0.02724 -0.0082 1.0000 0.1504
-3.000 -0.3395 0.03478 0.02499 -0.0074 1.0000 0.1506
-2.750 -0.3159 0.03299 0.02281 -0.0067 1.0000 0.1504
-2.500 -0.2912 0.03140 0.02085 -0.0062 1.0000 0.1537
-2.250 -0.2671 0.02993 0.01928 -0.0059 1.0000 0.1628
-2.000 -0.2413 0.02861 0.01772 -0.0055 1.0000 0.1717
-1.750 -0.2152 0.02742 0.01644 -0.0052 1.0000 0.1875
-1.500 -0.1874 0.02635 0.01534 -0.0051 1.0000 0.2105
-1.250 -0.1600 0.02529 0.01435 -0.0050 1.0000 0.2428
-1.000 -0.0974 0.02118 0.01356 -0.0096 1.0000 1.0000
-0.750 -0.0784 0.02157 0.01319 -0.0079 1.0000 1.0000
-0.500 -0.0617 0.02193 0.01312 -0.0067 1.0000 1.0000
-0.250 -0.0449 0.02231 0.01318 -0.0057 1.0000 1.0000
0.000 -0.0277 0.02273 0.01333 -0.0049 1.0000 1.0000
0.250 -0.0101 0.02318 0.01356 -0.0042 1.0000 1.0000
0.500 0.0077 0.02367 0.01385 -0.0036 1.0000 1.0000
0.750 0.0344 0.02432 0.01432 -0.0048 0.9967 1.0000
1.000 0.0734 0.02521 0.01503 -0.0084 0.9882 1.0000
1.250 0.1160 0.02620 0.01587 -0.0125 0.9772 1.0000
1.500 0.1756 0.02737 0.01692 -0.0196 0.9580 1.0000
1.750 0.2339 0.02833 0.01782 -0.0258 0.9333 1.0000
2.000 0.2978 0.02917 0.01865 -0.0325 0.9089 1.0000
2.250 0.3403 0.02982 0.01931 -0.0353 0.8873 1.0000
2.500 0.4071 0.03028 0.01986 -0.0415 0.8648 1.0000
2.750 0.4428 0.03073 0.02038 -0.0425 0.8417 1.0000
3.000 0.4982 0.03073 0.02051 -0.0454 0.8181 1.0000
3.250 0.5322 0.03077 0.02065 -0.0448 0.7926 1.0000
3.500 0.5648 0.03061 0.02059 -0.0433 0.7660 1.0000
3.750 0.5971 0.03005 0.02016 -0.0410 0.7380 1.0000
4.000 0.6253 0.02926 0.01946 -0.0375 0.7078 1.0000
4.250 0.6506 0.02818 0.01845 -0.0333 0.6733 1.0000
4.500 0.6771 0.02665 0.01695 -0.0287 0.6372 1.0000
4.750 0.7000 0.02537 0.01560 -0.0242 0.5928 1.0000
5.000 0.7231 0.02443 0.01446 -0.0204 0.5468 1.0000
5.250 0.7452 0.02419 0.01398 -0.0176 0.5031 1.0000
5.500 0.7675 0.02441 0.01398 -0.0154 0.4665 1.0000
5.750 0.7903 0.02499 0.01432 -0.0138 0.4370 1.0000
6.000 0.8137 0.02577 0.01487 -0.0125 0.4122 1.0000
6.250 0.8366 0.02687 0.01590 -0.0115 0.3908 1.0000
6.500 0.8600 0.02807 0.01708 -0.0107 0.3726 1.0000
6.750 0.8833 0.02937 0.01834 -0.0099 0.3560 1.0000
7.000 0.9061 0.03085 0.01987 -0.0092 0.3418 1.0000
7.250 0.9284 0.03245 0.02156 -0.0085 0.3287 1.0000
7.500 0.9502 0.03413 0.02335 -0.0078 0.3161 1.0000
7.750 0.9741 0.03586 0.02504 -0.0072 0.3051 1.0000
8.000 0.9904 0.03799 0.02755 -0.0064 0.2944 1.0000
8.250 1.0071 0.04024 0.03005 -0.0055 0.2842 1.0000
8.500 1.0306 0.04223 0.03203 -0.0049 0.2748 1.0000
8.750 1.0358 0.04539 0.03577 -0.0038 0.2668 1.0000
9.000 1.0543 0.04768 0.03812 -0.0030 0.2575 1.0000
9.250 1.0531 0.05158 0.04254 -0.0019 0.2515 1.0000
9.500 1.0731 0.05399 0.04497 -0.0011 0.2432 1.0000
9.750 1.0538 0.05934 0.05084 0.0000 0.2399 1.0000
10.000 1.0266 0.06551 0.05733 0.0004 0.2382 1.0000
10.250 0.9814 0.07364 0.06560 -0.0007 0.2406 1.0000
10.500 0.9402 0.08275 0.07469 -0.0039 0.2431 1.0000
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Polar data table (+)
Polar graphs
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