NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 50,000 Max Cl/Cd: 33.34 at α=6.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc10n1-il-50000-n5.txt Download as CSV file: xf-rc10n1-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-10(N)1 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.000 -0.5727 0.10547 0.09918 0.0046 1.0000 0.1376
-8.750 -0.5816 0.10119 0.09501 0.0005 1.0000 0.1443
-8.250 -0.5844 0.09287 0.08683 -0.0041 1.0000 0.1589
-7.750 -0.5820 0.07637 0.06991 -0.0181 1.0000 0.0672
-7.500 -0.5880 0.07118 0.06417 -0.0207 1.0000 0.0569
-7.250 -0.5752 0.06654 0.05962 -0.0207 1.0000 0.0549
-7.000 -0.5669 0.06245 0.05538 -0.0210 1.0000 0.0534
-6.750 -0.5577 0.05859 0.05128 -0.0210 1.0000 0.0530
-6.500 -0.5466 0.05499 0.04739 -0.0208 1.0000 0.0533
-6.250 -0.5338 0.05160 0.04367 -0.0203 1.0000 0.0536
-6.000 -0.5194 0.04836 0.04010 -0.0196 1.0000 0.0534
-5.750 -0.5042 0.04528 0.03665 -0.0187 1.0000 0.0526
-5.500 -0.4890 0.04248 0.03351 -0.0175 1.0000 0.0518
-5.250 -0.4752 0.04005 0.03074 -0.0157 1.0000 0.0513
-5.000 -0.4635 0.03799 0.02835 -0.0135 1.0000 0.0510
-4.750 -0.4518 0.03613 0.02618 -0.0111 1.0000 0.0508
-4.500 -0.4363 0.03432 0.02404 -0.0093 0.9988 0.0509
-4.250 -0.4007 0.03211 0.02138 -0.0109 0.9887 0.0521
-4.000 -0.3640 0.03037 0.01918 -0.0125 0.9797 0.0555
-3.750 -0.3249 0.02862 0.01706 -0.0144 0.9723 0.0579
-3.500 -0.2875 0.02683 0.01521 -0.0161 0.9649 0.0602
-3.250 -0.2486 0.02549 0.01375 -0.0179 0.9584 0.0640
-3.000 -0.2137 0.02454 0.01258 -0.0193 0.9499 0.0718
-2.750 -0.1792 0.02343 0.01134 -0.0209 0.9425 0.0799
-2.500 -0.1489 0.02255 0.01034 -0.0217 0.9333 0.0917
-2.250 -0.1194 0.02156 0.00943 -0.0225 0.9247 0.1226
-2.000 0.0210 0.01881 0.00962 -0.0370 0.9472 1.0000
-1.750 0.0514 0.01868 0.00918 -0.0388 0.9342 1.0000
-1.500 0.0799 0.01862 0.00884 -0.0400 0.9219 1.0000
-1.250 0.1068 0.01862 0.00861 -0.0408 0.9105 1.0000
-1.000 0.1332 0.01866 0.00845 -0.0412 0.8999 1.0000
-0.750 0.1569 0.01877 0.00840 -0.0411 0.8889 1.0000
-0.500 0.1800 0.01893 0.00839 -0.0409 0.8785 1.0000
-0.250 0.2036 0.01910 0.00843 -0.0406 0.8690 1.0000
0.000 0.2267 0.01930 0.00852 -0.0401 0.8596 1.0000
0.250 0.2487 0.01954 0.00868 -0.0395 0.8496 1.0000
0.500 0.2715 0.01979 0.00884 -0.0388 0.8406 1.0000
0.750 0.2940 0.02005 0.00904 -0.0381 0.8313 1.0000
1.000 0.3156 0.02036 0.00931 -0.0373 0.8214 1.0000
1.250 0.3381 0.02064 0.00957 -0.0365 0.8124 1.0000
1.500 0.3603 0.02094 0.00986 -0.0356 0.8029 1.0000
1.750 0.3817 0.02129 0.01021 -0.0346 0.7924 1.0000
2.000 0.4042 0.02160 0.01053 -0.0337 0.7827 1.0000
2.250 0.4272 0.02186 0.01082 -0.0327 0.7733 1.0000
2.500 0.4484 0.02223 0.01125 -0.0317 0.7619 1.0000
2.750 0.4700 0.02257 0.01165 -0.0306 0.7508 1.0000
3.000 0.4921 0.02286 0.01201 -0.0295 0.7400 1.0000
3.250 0.5151 0.02306 0.01228 -0.0281 0.7299 1.0000
3.500 0.5363 0.02337 0.01273 -0.0269 0.7172 1.0000
3.750 0.5577 0.02365 0.01313 -0.0256 0.7043 1.0000
4.000 0.5793 0.02388 0.01349 -0.0242 0.6910 1.0000
4.250 0.6011 0.02403 0.01377 -0.0227 0.6767 1.0000
4.500 0.6231 0.02400 0.01392 -0.0209 0.6608 1.0000
4.750 0.6440 0.02386 0.01391 -0.0187 0.6402 1.0000
5.000 0.6658 0.02344 0.01361 -0.0162 0.6172 1.0000
5.250 0.6863 0.02315 0.01345 -0.0138 0.5899 1.0000
5.500 0.7072 0.02286 0.01333 -0.0115 0.5597 1.0000
5.750 0.7270 0.02277 0.01336 -0.0094 0.5209 1.0000
6.000 0.7464 0.02271 0.01332 -0.0070 0.4717 1.0000
6.250 0.7638 0.02291 0.01329 -0.0044 0.4030 1.0000
6.500 0.7772 0.02380 0.01371 -0.0020 0.3207 1.0000
6.750 0.7874 0.02533 0.01479 0.0001 0.2467 1.0000
7.000 0.7971 0.02705 0.01612 0.0019 0.1945 1.0000
7.250 0.8081 0.02873 0.01756 0.0036 0.1624 1.0000
7.500 0.8201 0.03035 0.01903 0.0051 0.1385 1.0000
7.750 0.8338 0.03194 0.02055 0.0067 0.1227 1.0000
8.000 0.8495 0.03348 0.02210 0.0081 0.1090 1.0000
8.250 0.8659 0.03500 0.02370 0.0093 0.0965 1.0000
8.500 0.8880 0.03661 0.02555 0.0105 0.0873 1.0000
8.750 0.9087 0.03841 0.02746 0.0115 0.0793 1.0000
9.000 0.9262 0.04012 0.02926 0.0124 0.0718 1.0000
9.250 0.9486 0.04259 0.03202 0.0133 0.0671 1.0000
9.500 0.9654 0.04504 0.03482 0.0144 0.0625 1.0000
9.750 0.9800 0.04721 0.03697 0.0151 0.0581 1.0000
10.000 0.9879 0.05029 0.04065 0.0166 0.0551 1.0000
10.250 0.9932 0.05378 0.04461 0.0179 0.0533 1.0000
10.500 0.9935 0.05739 0.04861 0.0193 0.0519 1.0000
10.750 0.9888 0.06098 0.05254 0.0207 0.0509 1.0000
11.000 0.9790 0.06441 0.05623 0.0220 0.0500 1.0000
11.250 0.9670 0.06809 0.06011 0.0226 0.0492 1.0000
11.500 0.9526 0.07224 0.06446 0.0223 0.0487 1.0000
11.750 0.9288 0.07782 0.07027 0.0206 0.0490 1.0000
12.000 0.8950 0.08565 0.07832 0.0165 0.0504 1.0000
12.250 0.8628 0.09473 0.08752 0.0108 0.0518 1.0000
|
Polar data table (+)
Polar graphs
<< Back to NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il)