NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-10(N)1 AIRFOIL (rc10n1-il) Reynolds number: 100,000 Max Cl/Cd: 46.82 at α=5.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rc10n1-il-100000-n5.txt Download as CSV file: xf-rc10n1-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5779 0.09232 0.08788 -0.0024 1.0000 0.0481
-8.500 -0.5804 0.08701 0.08262 -0.0073 1.0000 0.0484
-8.250 -0.5853 0.08200 0.07761 -0.0114 1.0000 0.0487
-8.000 -0.5880 0.07742 0.07298 -0.0141 1.0000 0.0490
-7.750 -0.5878 0.07289 0.06837 -0.0163 1.0000 0.0494
-7.500 -0.5847 0.06866 0.06405 -0.0178 1.0000 0.0506
-7.000 -0.5711 0.05866 0.05350 -0.0198 1.0000 0.0412
-6.750 -0.5610 0.05438 0.04906 -0.0199 1.0000 0.0392
-6.250 -0.5325 0.04527 0.03897 -0.0183 1.0000 0.0306
-6.000 -0.5141 0.04158 0.03498 -0.0185 0.9804 0.0301
-5.750 -0.4877 0.03796 0.03095 -0.0195 0.9567 0.0298
-5.500 -0.4614 0.03485 0.02735 -0.0200 0.9396 0.0296
-5.250 -0.4349 0.03249 0.02448 -0.0198 0.9255 0.0304
-5.000 -0.4096 0.03059 0.02207 -0.0192 0.9131 0.0319
-4.750 -0.3854 0.02823 0.01936 -0.0186 0.9021 0.0323
-4.500 -0.3601 0.02636 0.01717 -0.0180 0.8919 0.0325
-4.250 -0.3345 0.02466 0.01522 -0.0173 0.8831 0.0329
-4.000 -0.3086 0.02319 0.01353 -0.0167 0.8745 0.0335
-3.750 -0.2826 0.02193 0.01215 -0.0162 0.8665 0.0344
-3.500 -0.2572 0.02086 0.01098 -0.0155 0.8591 0.0356
-3.250 -0.2318 0.02000 0.01005 -0.0149 0.8513 0.0382
-3.000 -0.2072 0.01932 0.00926 -0.0141 0.8445 0.0419
-2.750 -0.1829 0.01858 0.00844 -0.0133 0.8371 0.0442
-2.500 -0.1610 0.01775 0.00757 -0.0121 0.8306 0.0477
-2.250 -0.1366 0.01728 0.00700 -0.0115 0.8236 0.0547
-2.000 -0.1126 0.01676 0.00641 -0.0107 0.8173 0.0658
-1.750 -0.0881 0.01615 0.00585 -0.0100 0.8107 0.0961
-1.500 -0.0783 0.01331 0.00572 -0.0065 0.8048 0.7282
-1.250 -0.0190 0.01354 0.00632 -0.0095 0.8015 0.9314
-1.000 0.0410 0.01380 0.00635 -0.0154 0.7961 0.9692
-0.750 0.0798 0.01379 0.00615 -0.0179 0.7901 0.9801
-0.500 0.1189 0.01377 0.00597 -0.0205 0.7842 0.9905
-0.250 0.1592 0.01373 0.00580 -0.0236 0.7776 1.0000
0.000 0.1830 0.01374 0.00569 -0.0231 0.7714 1.0000
0.250 0.2078 0.01377 0.00564 -0.0229 0.7638 1.0000
0.500 0.2317 0.01381 0.00559 -0.0223 0.7577 1.0000
0.750 0.2568 0.01387 0.00561 -0.0222 0.7498 1.0000
1.000 0.2809 0.01393 0.00559 -0.0215 0.7435 1.0000
1.250 0.3058 0.01401 0.00566 -0.0212 0.7350 1.0000
1.500 0.3301 0.01407 0.00569 -0.0206 0.7278 1.0000
1.750 0.3548 0.01415 0.00576 -0.0201 0.7194 1.0000
2.000 0.3794 0.01424 0.00584 -0.0196 0.7113 1.0000
2.250 0.4038 0.01430 0.00590 -0.0189 0.7033 1.0000
2.500 0.4285 0.01440 0.00605 -0.0184 0.6939 1.0000
2.750 0.4528 0.01445 0.00610 -0.0176 0.6859 1.0000
3.000 0.4775 0.01454 0.00624 -0.0171 0.6760 1.0000
3.250 0.5022 0.01463 0.00641 -0.0165 0.6659 1.0000
3.500 0.5266 0.01468 0.00651 -0.0157 0.6557 1.0000
3.750 0.5507 0.01468 0.00656 -0.0148 0.6423 1.0000
4.000 0.5744 0.01463 0.00656 -0.0138 0.6229 1.0000
4.250 0.5979 0.01454 0.00649 -0.0125 0.6020 1.0000
4.500 0.6218 0.01457 0.00658 -0.0116 0.5782 1.0000
4.750 0.6455 0.01462 0.00667 -0.0106 0.5527 1.0000
5.000 0.6688 0.01472 0.00677 -0.0096 0.5214 1.0000
5.250 0.6915 0.01489 0.00689 -0.0085 0.4796 1.0000
5.500 0.7130 0.01523 0.00710 -0.0072 0.4241 1.0000
5.750 0.7325 0.01588 0.00742 -0.0059 0.3547 1.0000
6.000 0.7503 0.01684 0.00800 -0.0047 0.2776 1.0000
6.250 0.7671 0.01798 0.00875 -0.0036 0.2044 1.0000
6.500 0.7843 0.01911 0.00958 -0.0025 0.1538 1.0000
6.750 0.8022 0.02015 0.01051 -0.0014 0.1249 1.0000
7.000 0.8204 0.02112 0.01144 -0.0003 0.1048 1.0000
7.250 0.8384 0.02210 0.01240 0.0008 0.0889 1.0000
7.500 0.8558 0.02315 0.01346 0.0020 0.0784 1.0000
7.750 0.8722 0.02428 0.01457 0.0033 0.0696 1.0000
8.000 0.8902 0.02524 0.01568 0.0044 0.0612 1.0000
8.250 0.9056 0.02654 0.01700 0.0058 0.0562 1.0000
8.500 0.9230 0.02769 0.01830 0.0070 0.0509 1.0000
8.750 0.9378 0.02903 0.01962 0.0082 0.0463 1.0000
9.000 0.9555 0.03040 0.02119 0.0094 0.0428 1.0000
9.250 0.9727 0.03195 0.02291 0.0106 0.0399 1.0000
9.500 0.9885 0.03343 0.02451 0.0117 0.0372 1.0000
9.750 1.0029 0.03536 0.02650 0.0127 0.0345 1.0000
10.000 1.0183 0.03721 0.02870 0.0139 0.0323 1.0000
10.250 1.0317 0.03950 0.03129 0.0152 0.0306 1.0000
10.500 1.0422 0.04178 0.03385 0.0165 0.0291 1.0000
10.750 1.0497 0.04374 0.03599 0.0179 0.0277 1.0000
11.000 1.0542 0.04575 0.03808 0.0193 0.0264 1.0000
11.250 1.0520 0.04869 0.04125 0.0210 0.0254 1.0000
11.500 1.0459 0.05166 0.04459 0.0225 0.0247 1.0000
11.750 1.0366 0.05523 0.04850 0.0234 0.0243 1.0000
12.000 1.0240 0.05926 0.05284 0.0236 0.0240 1.0000
12.250 1.0086 0.06379 0.05764 0.0230 0.0239 1.0000
12.500 0.9908 0.06890 0.06301 0.0216 0.0238 1.0000
12.750 0.9710 0.07469 0.06902 0.0192 0.0238 1.0000
13.000 0.9492 0.08134 0.07588 0.0157 0.0239 1.0000
13.250 0.9259 0.08902 0.08372 0.0110 0.0241 1.0000
13.500 0.9011 0.09799 0.09283 0.0052 0.0244 1.0000
13.750 0.8753 0.10845 0.10337 -0.0015 0.0248 1.0000
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Polar data table (+)
Polar graphs
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