NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: NASA/LANGLEY RC-SC2 AIRFOIL (rcsc2-il) Reynolds number: 100,000 Max Cl/Cd: 33.68 at α=2.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rcsc2-il-100000-n5.txt Download as CSV file: xf-rcsc2-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: NASA/LANGLEY RC-SC2 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
------ -------- --------- --------- -------- -------- --------
-10.750 -0.6342 0.08983 0.08391 -0.0341 1.0000 0.0382
-10.500 -0.6476 0.08149 0.07557 -0.0398 1.0000 0.0379
-10.250 -0.6679 0.07418 0.06820 -0.0448 1.0000 0.0377
-10.000 -0.6920 0.06859 0.06251 -0.0468 1.0000 0.0375
-9.750 -0.7174 0.06447 0.05828 -0.0454 1.0000 0.0373
-9.500 -0.7410 0.06082 0.05447 -0.0423 1.0000 0.0373
-9.250 -0.7600 0.05706 0.05049 -0.0390 1.0000 0.0373
-9.000 -0.7740 0.05345 0.04660 -0.0356 1.0000 0.0374
-8.750 -0.7827 0.05005 0.04289 -0.0321 1.0000 0.0375
-8.500 -0.7867 0.04687 0.03936 -0.0287 1.0000 0.0376
-8.250 -0.7864 0.04390 0.03605 -0.0255 1.0000 0.0376
-8.000 -0.7827 0.04112 0.03289 -0.0225 1.0000 0.0377
-7.750 -0.7759 0.03855 0.02993 -0.0197 1.0000 0.0379
-7.500 -0.7660 0.03620 0.02718 -0.0171 1.0000 0.0381
-7.250 -0.7525 0.03427 0.02505 -0.0151 1.0000 0.0385
-7.000 -0.7370 0.03266 0.02326 -0.0133 1.0000 0.0389
-6.750 -0.7203 0.03119 0.02164 -0.0117 1.0000 0.0394
-6.500 -0.7026 0.02984 0.02013 -0.0101 1.0000 0.0401
-6.250 -0.6841 0.02868 0.01882 -0.0086 1.0000 0.0413
-6.000 -0.6647 0.02753 0.01747 -0.0071 1.0000 0.0428
-5.750 -0.6444 0.02636 0.01608 -0.0057 1.0000 0.0442
-5.500 -0.6237 0.02522 0.01482 -0.0044 1.0000 0.0453
-5.250 -0.6032 0.02424 0.01386 -0.0032 1.0000 0.0464
-5.000 -0.5824 0.02339 0.01297 -0.0020 1.0000 0.0479
-4.750 -0.5614 0.02258 0.01211 -0.0007 1.0000 0.0497
-4.500 -0.5402 0.02187 0.01131 0.0005 1.0000 0.0523
-4.250 -0.5196 0.02116 0.01065 0.0017 1.0000 0.0553
-4.000 -0.4984 0.02055 0.01004 0.0029 1.0000 0.0591
-3.750 -0.4770 0.01993 0.00937 0.0041 1.0000 0.0632
-3.500 -0.4556 0.01936 0.00886 0.0051 1.0000 0.0696
-3.250 -0.4339 0.01880 0.00834 0.0061 1.0000 0.0790
-3.000 -0.4119 0.01827 0.00788 0.0071 1.0000 0.0933
-2.750 -0.3896 0.01775 0.00752 0.0079 1.0000 0.1173
-2.500 -0.3672 0.01724 0.00723 0.0087 1.0000 0.1581
-2.250 -0.3451 0.01654 0.00700 0.0093 1.0000 0.2419
-2.000 -0.3247 0.01527 0.00685 0.0101 1.0000 0.4527
-1.750 -0.3055 0.01445 0.00710 0.0126 1.0000 0.6827
-1.500 -0.2827 0.01445 0.00744 0.0148 0.9998 0.8010
-1.250 -0.2464 0.01472 0.00779 0.0141 0.9971 0.8782
-1.000 -0.1991 0.01516 0.00820 0.0110 0.9963 0.9349
-0.750 -0.1413 0.01560 0.00856 0.0054 0.9974 0.9745
-0.500 -0.0662 0.01596 0.00881 -0.0042 1.0000 1.0000
-0.250 -0.0355 0.01599 0.00876 -0.0055 0.9956 1.0000
0.000 0.0010 0.01604 0.00876 -0.0079 0.9896 1.0000
0.250 0.0399 0.01613 0.00882 -0.0107 0.9839 1.0000
0.500 0.0786 0.01616 0.00884 -0.0133 0.9758 1.0000
0.750 0.1300 0.01612 0.00880 -0.0181 0.9645 1.0000
1.000 0.1939 0.01587 0.00860 -0.0251 0.9489 1.0000
1.250 0.2463 0.01557 0.00835 -0.0297 0.9338 1.0000
1.500 0.2864 0.01528 0.00812 -0.0318 0.9181 1.0000
1.750 0.3191 0.01494 0.00784 -0.0322 0.8968 1.0000
2.000 0.3538 0.01451 0.00747 -0.0328 0.8689 1.0000
2.250 0.3846 0.01411 0.00711 -0.0325 0.8262 1.0000
2.500 0.4321 0.01359 0.00640 -0.0347 0.7345 1.0000
2.750 0.4655 0.01382 0.00591 -0.0344 0.5724 1.0000
3.000 0.4822 0.01468 0.00600 -0.0320 0.4207 1.0000
3.250 0.4986 0.01549 0.00626 -0.0300 0.3215 1.0000
3.500 0.5165 0.01615 0.00657 -0.0283 0.2560 1.0000
3.750 0.5347 0.01676 0.00690 -0.0266 0.2020 1.0000
4.000 0.5527 0.01739 0.00724 -0.0249 0.1596 1.0000
4.250 0.5713 0.01795 0.00762 -0.0233 0.1298 1.0000
4.500 0.5900 0.01851 0.00806 -0.0217 0.1095 1.0000
4.750 0.6086 0.01909 0.00856 -0.0200 0.0935 1.0000
5.000 0.6270 0.01972 0.00911 -0.0182 0.0812 1.0000
5.250 0.6450 0.02040 0.00973 -0.0165 0.0718 1.0000
5.500 0.6634 0.02107 0.01040 -0.0147 0.0648 1.0000
5.750 0.6814 0.02185 0.01118 -0.0129 0.0598 1.0000
6.000 0.6999 0.02256 0.01188 -0.0112 0.0554 1.0000
6.250 0.7181 0.02338 0.01270 -0.0095 0.0522 1.0000
6.500 0.7374 0.02422 0.01363 -0.0079 0.0494 1.0000
6.750 0.7566 0.02513 0.01457 -0.0064 0.0475 1.0000
7.000 0.7757 0.02609 0.01551 -0.0049 0.0460 1.0000
7.250 0.7954 0.02725 0.01669 -0.0036 0.0448 1.0000
7.500 0.8161 0.02851 0.01813 -0.0023 0.0437 1.0000
7.750 0.8363 0.02983 0.01963 -0.0010 0.0425 1.0000
8.000 0.8555 0.03113 0.02110 0.0004 0.0413 1.0000
8.250 0.8739 0.03234 0.02244 0.0017 0.0401 1.0000
8.500 0.8916 0.03353 0.02370 0.0031 0.0391 1.0000
8.750 0.9090 0.03495 0.02516 0.0044 0.0383 1.0000
9.000 0.9235 0.03695 0.02751 0.0062 0.0378 1.0000
9.250 0.9355 0.03920 0.03012 0.0083 0.0374 1.0000
9.500 0.9445 0.04168 0.03297 0.0106 0.0371 1.0000
9.750 0.9498 0.04439 0.03605 0.0131 0.0369 1.0000
10.000 0.9510 0.04734 0.03938 0.0159 0.0367 1.0000
10.250 0.9473 0.05053 0.04293 0.0188 0.0366 1.0000
10.500 0.9384 0.05395 0.04668 0.0219 0.0365 1.0000
10.750 0.9225 0.05738 0.05040 0.0254 0.0366 1.0000
11.000 0.9016 0.06118 0.05445 0.0282 0.0366 1.0000
11.250 0.8766 0.06578 0.05928 0.0295 0.0368 1.0000
11.500 0.8471 0.07175 0.06548 0.0282 0.0370 1.0000
11.750 0.8117 0.08058 0.07450 0.0227 0.0373 1.0000
12.000 0.7736 0.09436 0.08840 0.0122 0.0378 1.0000
12.250 0.7550 0.10452 0.09856 0.0064 0.0382 1.0000
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Polar data table (+)
Polar graphs
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