Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NASA SC(2)-0710 AIRFOIL (sc20710-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NASA SC(2)-0710 AIRFOIL (sc20710-il)
Reynolds number: 200,000
Max Cl/Cd: 45.02 at α=0.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-sc20710-il-200000-n5.txt
Download as CSV file: xf-sc20710-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.750  -0.7054   0.08055   0.07637  -0.0449   1.0000   0.0230
 -12.500  -0.7827   0.06129   0.05681  -0.0588   1.0000   0.0222
 -12.250  -0.8175   0.05471   0.05001  -0.0611   1.0000   0.0221
 -12.000  -0.8456   0.05042   0.04554  -0.0603   1.0000   0.0221
 -11.750  -0.8713   0.04740   0.04237  -0.0575   1.0000   0.0222
 -11.500  -0.8860   0.04383   0.03855  -0.0568   1.0000   0.0224
 -11.250  -0.8913   0.04034   0.03472  -0.0566   1.0000   0.0227
 -11.000  -0.8890   0.03708   0.03108  -0.0563   1.0000   0.0231
 -10.750  -0.8800   0.03419   0.02779  -0.0561   1.0000   0.0234
 -10.500  -0.8660   0.03175   0.02497  -0.0560   1.0000   0.0239
 -10.250  -0.8489   0.02985   0.02284  -0.0559   1.0000   0.0243
 -10.000  -0.8295   0.02853   0.02143  -0.0557   1.0000   0.0248
  -9.750  -0.8088   0.02739   0.02019  -0.0557   1.0000   0.0254
  -9.500  -0.7871   0.02628   0.01895  -0.0556   1.0000   0.0261
  -9.250  -0.7647   0.02516   0.01769  -0.0556   1.0000   0.0270
  -9.000  -0.7414   0.02413   0.01648  -0.0556   1.0000   0.0283
  -8.750  -0.7175   0.02314   0.01530  -0.0556   1.0000   0.0296
  -8.500  -0.6933   0.02226   0.01444  -0.0559   1.0000   0.0308
  -8.250  -0.6683   0.02152   0.01365  -0.0561   1.0000   0.0321
  -8.000  -0.6428   0.02076   0.01279  -0.0563   1.0000   0.0338
  -7.750  -0.6166   0.02005   0.01196  -0.0565   1.0000   0.0355
  -7.500  -0.5893   0.01930   0.01123  -0.0573   1.0000   0.0375
  -7.250  -0.5620   0.01884   0.01074  -0.0578   1.0000   0.0402
  -7.000  -0.5313   0.01837   0.01017  -0.0589   0.9993   0.0432
  -6.750  -0.4986   0.01775   0.00960  -0.0607   0.9985   0.0466
  -6.500  -0.4660   0.01735   0.00914  -0.0622   0.9975   0.0507
  -6.250  -0.4326   0.01683   0.00864  -0.0640   0.9967   0.0551
  -6.000  -0.3994   0.01646   0.00826  -0.0657   0.9958   0.0604
  -5.750  -0.3651   0.01602   0.00781  -0.0676   0.9951   0.0661
  -5.500  -0.3307   0.01568   0.00746  -0.0695   0.9943   0.0739
  -5.250  -0.2962   0.01528   0.00713  -0.0714   0.9937   0.0835
  -5.000  -0.2626   0.01495   0.00683  -0.0732   0.9928   0.0981
  -4.750  -0.2283   0.01457   0.00657  -0.0751   0.9920   0.1215
  -4.500  -0.1927   0.01414   0.00635  -0.0775   0.9915   0.1661
  -4.250  -0.1535   0.01348   0.00611  -0.0810   0.9916   0.2595
  -4.000  -0.1086   0.01254   0.00596  -0.0860   0.9931   0.4364
  -3.750  -0.0759   0.01239   0.00638  -0.0871   0.9921   0.5687
  -3.500  -0.0461   0.01253   0.00659  -0.0873   0.9901   0.6064
  -3.250  -0.0148   0.01268   0.00674  -0.0880   0.9882   0.6299
  -3.000   0.0159   0.01290   0.00698  -0.0884   0.9863   0.6499
  -2.750   0.0468   0.01318   0.00727  -0.0888   0.9845   0.6675
  -2.500   0.0758   0.01343   0.00754  -0.0888   0.9817   0.6816
  -2.250   0.1046   0.01362   0.00774  -0.0888   0.9783   0.6910
  -2.000   0.1345   0.01381   0.00795  -0.0891   0.9756   0.6983
  -1.750   0.1663   0.01398   0.00814  -0.0897   0.9734   0.7054
  -1.500   0.2021   0.01405   0.00821  -0.0914   0.9713   0.7117
  -1.250   0.2284   0.01413   0.00834  -0.0908   0.9655   0.7152
  -1.000   0.2657   0.01406   0.00830  -0.0926   0.9608   0.7190
  -0.750   0.3041   0.01386   0.00814  -0.0946   0.9539   0.7230
  -0.500   0.3579   0.01326   0.00754  -0.0995   0.9442   0.7266
  -0.250   0.4028   0.01256   0.00690  -0.1019   0.9267   0.7281
   0.250   0.4740   0.01181   0.00622  -0.1035   0.8784   0.7318
   0.500   0.5127   0.01152   0.00592  -0.1050   0.8347   0.7342
   0.750   0.5461   0.01213   0.00542  -0.1047   0.5938   0.7366
   1.000   0.5619   0.01347   0.00579  -0.1025   0.3996   0.7397
   1.500   0.6063   0.01520   0.00647  -0.1009   0.1790   0.7446
   1.750   0.6302   0.01577   0.00681  -0.1002   0.1289   0.7467
   2.000   0.6555   0.01623   0.00715  -0.0997   0.1035   0.7490
   2.250   0.6818   0.01664   0.00751  -0.0994   0.0875   0.7514
   2.500   0.7088   0.01704   0.00787  -0.0993   0.0767   0.7540
   2.750   0.7362   0.01751   0.00827  -0.0994   0.0689   0.7568
   3.000   0.7634   0.01788   0.00868  -0.0993   0.0629   0.7591
   3.250   0.7878   0.01842   0.00919  -0.0986   0.0577   0.7607
   3.500   0.8134   0.01882   0.00967  -0.0981   0.0535   0.7625
   3.750   0.8386   0.01929   0.01013  -0.0976   0.0494   0.7646
   4.000   0.8637   0.01984   0.01073  -0.0971   0.0460   0.7672
   4.250   0.8898   0.02033   0.01126  -0.0968   0.0427   0.7700
   4.500   0.9159   0.02095   0.01187  -0.0966   0.0402   0.7732
   4.750   0.9405   0.02163   0.01262  -0.0959   0.0378   0.7754
   5.000   0.9647   0.02221   0.01329  -0.0952   0.0354   0.7773
   5.250   0.9888   0.02280   0.01390  -0.0946   0.0334   0.7793
   5.500   1.0126   0.02368   0.01484  -0.0938   0.0318   0.7815
   5.750   1.0371   0.02456   0.01585  -0.0932   0.0303   0.7838
   6.000   1.0618   0.02546   0.01686  -0.0927   0.0290   0.7864
   6.250   1.0867   0.02637   0.01783  -0.0923   0.0280   0.7892
   6.500   1.1094   0.02733   0.01884  -0.0915   0.0271   0.7911
   6.750   1.1315   0.02859   0.02025  -0.0905   0.0263   0.7929
   7.000   1.1539   0.02985   0.02177  -0.0895   0.0254   0.7951
   7.250   1.1756   0.03111   0.02326  -0.0885   0.0243   0.7976
   7.500   1.1969   0.03232   0.02466  -0.0875   0.0235   0.8005
   7.750   1.2177   0.03366   0.02617  -0.0866   0.0230   0.8034
   8.000   1.2372   0.03497   0.02763  -0.0855   0.0225   0.8059
   8.250   1.2545   0.03639   0.02920  -0.0840   0.0221   0.8078
   8.500   1.2694   0.03832   0.03133  -0.0822   0.0218   0.8099
   8.750   1.2795   0.04108   0.03458  -0.0798   0.0214   0.8124
   9.000   1.2841   0.04443   0.03844  -0.0768   0.0211   0.8151
   9.250   1.2816   0.04845   0.04299  -0.0732   0.0208   0.8178
   9.500   1.2708   0.05301   0.04806  -0.0691   0.0206   0.8204
   9.750   1.2491   0.05758   0.05307  -0.0640   0.0205   0.8222
  10.000   1.2203   0.06218   0.05803  -0.0588   0.0204   0.8243
  10.250   1.1868   0.06799   0.06419  -0.0556   0.0204   0.8263
  10.500   1.1477   0.07549   0.07201  -0.0554   0.0205   0.8282
  10.750   1.1027   0.08638   0.08319  -0.0609   0.0207   0.8298
<< Back to NASA SC(2)-0710 AIRFOIL (sc20710-il)

Polar data table (+)

Polar graphs


<< Back to NASA SC(2)-0710 AIRFOIL (sc20710-il)