NASA SC(2)-0710 AIRFOIL (sc20710-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: NASA SC(2)-0710 AIRFOIL (sc20710-il) Reynolds number: 50,000 Max Cl/Cd: 24.76 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-sc20710-il-50000.txt Download as CSV file: xf-sc20710-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: NASA SC(2)-0710 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.4717 0.09982 0.09206 -0.0047 1.0000 0.3929
-8.750 -0.4762 0.09730 0.08962 -0.0036 1.0000 0.4052
-8.500 -0.4849 0.09537 0.08778 -0.0020 1.0000 0.4190
-8.250 -0.6554 0.06248 0.05475 -0.0444 1.0000 0.1595
-8.000 -0.6519 0.05428 0.04570 -0.0509 1.0000 0.1469
-7.750 -0.6344 0.04960 0.04064 -0.0527 1.0000 0.1460
-7.500 -0.6112 0.04513 0.03554 -0.0555 1.0000 0.1470
-7.250 -0.5837 0.04119 0.03095 -0.0578 1.0000 0.1480
-7.000 -0.5600 0.03805 0.02776 -0.0581 1.0000 0.1522
-6.750 -0.5327 0.03554 0.02491 -0.0589 1.0000 0.1593
-6.500 -0.5027 0.03292 0.02179 -0.0599 1.0000 0.1652
-6.250 -0.4773 0.03105 0.01987 -0.0597 1.0000 0.1758
-6.000 -0.4515 0.02919 0.01789 -0.0591 1.0000 0.1854
-5.750 -0.4271 0.02771 0.01638 -0.0581 1.0000 0.2002
-5.500 -0.4045 0.02640 0.01516 -0.0564 1.0000 0.2176
-5.250 -0.3817 0.02514 0.01404 -0.0548 1.0000 0.2444
-5.000 -0.3579 0.02356 0.01293 -0.0539 1.0000 0.2930
-4.750 -0.3460 0.02178 0.01345 -0.0488 1.0000 0.5014
-4.500 -0.3540 0.02478 0.01663 -0.0351 1.0000 0.6629
-4.250 -0.3568 0.02662 0.01835 -0.0244 1.0000 0.7119
-4.000 -0.3579 0.02763 0.01925 -0.0148 1.0000 0.7491
-3.750 -0.3559 0.02804 0.01954 -0.0067 1.0000 0.7859
-3.500 -0.3525 0.02794 0.01932 0.0005 1.0000 0.8219
-3.250 -0.3489 0.02731 0.01858 0.0077 1.0000 0.8576
-3.000 -0.3405 0.02633 0.01745 0.0133 1.0000 0.8959
-2.750 -0.1385 0.02490 0.01515 -0.0155 1.0000 0.9859
-2.500 -0.0777 0.02422 0.01426 -0.0235 1.0000 1.0000
-2.250 -0.0710 0.02384 0.01383 -0.0212 1.0000 1.0000
-2.000 -0.0643 0.02348 0.01345 -0.0188 1.0000 1.0000
-1.750 -0.0578 0.02315 0.01310 -0.0163 1.0000 1.0000
-1.500 -0.0513 0.02284 0.01278 -0.0138 1.0000 1.0000
-1.250 -0.0450 0.02255 0.01249 -0.0113 1.0000 1.0000
-1.000 -0.0387 0.02227 0.01221 -0.0087 1.0000 1.0000
-0.750 -0.0323 0.02200 0.01196 -0.0062 1.0000 1.0000
-0.500 -0.0257 0.02175 0.01173 -0.0037 1.0000 1.0000
-0.250 -0.0184 0.02152 0.01153 -0.0014 1.0000 1.0000
0.000 -0.0080 0.02136 0.01139 0.0004 1.0000 1.0000
0.250 0.0078 0.02130 0.01137 0.0011 1.0000 1.0000
0.500 0.0280 0.02136 0.01147 0.0010 1.0000 1.0000
0.750 0.0516 0.02151 0.01168 0.0003 1.0000 1.0000
1.000 0.0774 0.02176 0.01200 -0.0009 1.0000 1.0000
1.250 0.1047 0.02208 0.01241 -0.0024 1.0000 1.0000
1.500 0.1329 0.02248 0.01292 -0.0040 1.0000 1.0000
1.750 0.1617 0.02295 0.01352 -0.0059 1.0000 1.0000
2.000 0.1906 0.02350 0.01423 -0.0078 1.0000 1.0000
2.250 0.2195 0.02413 0.01504 -0.0099 1.0000 1.0000
2.500 0.2481 0.02486 0.01597 -0.0120 1.0000 1.0000
2.750 0.2762 0.02570 0.01706 -0.0142 1.0000 1.0000
3.000 0.3034 0.02669 0.01830 -0.0165 1.0000 1.0000
3.250 0.4468 0.02624 0.01865 -0.0346 0.9169 1.0000
3.500 0.5632 0.02275 0.01254 -0.0321 0.3041 1.0000
3.750 0.6022 0.02503 0.01421 -0.0340 0.2375 1.0000
4.000 0.6488 0.02714 0.01600 -0.0372 0.2036 1.0000
4.250 0.6888 0.02921 0.01813 -0.0391 0.1851 1.0000
4.500 0.7245 0.03147 0.02039 -0.0406 0.1714 1.0000
4.750 0.7562 0.03355 0.02289 -0.0411 0.1615 1.0000
5.000 0.7868 0.03623 0.02579 -0.0419 0.1539 1.0000
5.250 0.8144 0.03898 0.02910 -0.0419 0.1488 1.0000
5.500 0.8426 0.04188 0.03205 -0.0426 0.1424 1.0000
5.750 0.8675 0.04578 0.03629 -0.0427 0.1411 1.0000
6.000 0.8878 0.04951 0.04060 -0.0422 0.1407 1.0000
6.250 0.9047 0.05349 0.04512 -0.0415 0.1405 1.0000
6.500 0.9093 0.05831 0.05098 -0.0399 0.1468 1.0000
6.750 0.9205 0.06404 0.05711 -0.0399 0.1542 1.0000
7.000 0.9354 0.07170 0.06509 -0.0409 0.1720 1.0000
7.250 0.7395 0.10349 0.09811 -0.0825 0.4320 1.0000
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Polar data table (+)
Polar graphs
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