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NASA SC(2)-0710 AIRFOIL (sc20710-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0710 AIRFOIL (sc20710-il)
Reynolds number: 100,000
Max Cl/Cd: 32.33 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20710-il-100000.txt
Download as CSV file: xf-sc20710-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5548   0.10018   0.09471  -0.0281   1.0000   0.1510
 -10.000  -0.5415   0.09774   0.09227  -0.0262   1.0000   0.1563
  -9.750  -0.5938   0.09314   0.08787  -0.0319   1.0000   0.1632
  -9.500  -0.5674   0.08982   0.08453  -0.0284   1.0000   0.1672
  -9.250  -0.5599   0.08723   0.08195  -0.0267   1.0000   0.1720
  -9.000  -0.6431   0.08106   0.07599  -0.0349   1.0000   0.1792
  -8.750  -0.6085   0.07923   0.07420  -0.0275   1.0000   0.1833
  -8.500  -0.6886   0.04997   0.04340  -0.0526   1.0000   0.0911
  -8.250  -0.6715   0.04493   0.03796  -0.0537   1.0000   0.0849
  -8.000  -0.6507   0.03986   0.03236  -0.0555   1.0000   0.0819
  -7.750  -0.6252   0.03569   0.02761  -0.0572   1.0000   0.0810
  -7.500  -0.5988   0.03308   0.02458  -0.0582   1.0000   0.0836
  -7.250  -0.5702   0.03078   0.02177  -0.0592   1.0000   0.0871
  -7.000  -0.5415   0.02845   0.01901  -0.0599   1.0000   0.0896
  -6.750  -0.5165   0.02672   0.01734  -0.0600   1.0000   0.0950
  -6.500  -0.4888   0.02547   0.01581  -0.0601   1.0000   0.1012
  -6.250  -0.4627   0.02389   0.01424  -0.0601   1.0000   0.1076
  -6.000  -0.4356   0.02291   0.01316  -0.0600   1.0000   0.1157
  -5.750  -0.4094   0.02171   0.01208  -0.0598   1.0000   0.1244
  -5.500  -0.3818   0.02074   0.01112  -0.0598   1.0000   0.1358
  -5.250  -0.3533   0.01981   0.01035  -0.0601   1.0000   0.1513
  -5.000  -0.3225   0.01887   0.00962  -0.0611   1.0000   0.1788
  -4.750  -0.2717   0.01605   0.00914  -0.0675   1.0000   0.4973
  -4.500  -0.2619   0.01718   0.01056  -0.0616   1.0000   0.6322
  -4.250  -0.2516   0.01826   0.01160  -0.0561   1.0000   0.6668
  -4.000  -0.2383   0.01922   0.01248  -0.0517   1.0000   0.6942
  -3.750  -0.2293   0.02003   0.01326  -0.0462   1.0000   0.7119
  -3.500  -0.2187   0.02068   0.01387  -0.0412   1.0000   0.7290
  -3.250  -0.2080   0.02119   0.01435  -0.0364   1.0000   0.7458
  -3.000  -0.1983   0.02158   0.01472  -0.0314   1.0000   0.7628
  -2.750  -0.1886   0.02184   0.01496  -0.0266   1.0000   0.7806
  -2.500  -0.1786   0.02198   0.01508  -0.0220   1.0000   0.7995
  -2.250  -0.1652   0.02206   0.01512  -0.0183   1.0000   0.8209
  -2.000  -0.1584   0.02181   0.01487  -0.0131   1.0000   0.8390
  -1.750  -0.1463   0.02151   0.01455  -0.0096   1.0000   0.8560
  -1.500  -0.1289   0.02119   0.01421  -0.0075   1.0000   0.8700
  -1.250  -0.1066   0.02096   0.01393  -0.0069   1.0000   0.8815
  -1.000  -0.0844   0.02070   0.01366  -0.0063   1.0000   0.8912
  -0.750  -0.0620   0.02042   0.01337  -0.0058   1.0000   0.8992
  -0.500  -0.0340   0.02036   0.01330  -0.0067   1.0000   0.9062
  -0.250  -0.0106   0.02015   0.01309  -0.0066   1.0000   0.9125
   0.000   0.0166   0.02012   0.01308  -0.0074   1.0000   0.9184
   0.250   0.0438   0.02010   0.01310  -0.0083   1.0000   0.9234
   0.500   0.0695   0.02009   0.01312  -0.0088   1.0000   0.9285
   0.750   0.0987   0.02024   0.01333  -0.0102   1.0000   0.9329
   1.000   0.1264   0.02038   0.01354  -0.0113   1.0000   0.9370
   1.250   0.1534   0.02056   0.01380  -0.0124   1.0000   0.9410
   1.500   0.1816   0.02085   0.01419  -0.0138   1.0000   0.9453
   1.750   0.2276   0.02121   0.01469  -0.0184   0.9900   0.9486
   2.000   0.3128   0.02089   0.01456  -0.0287   0.9556   0.9500
   2.250   0.3775   0.01986   0.01374  -0.0342   0.9249   0.9523
   2.500   0.4327   0.01845   0.01255  -0.0372   0.8934   0.9549
   2.750   0.4792   0.01656   0.01090  -0.0375   0.8476   0.9579
   3.000   0.5386   0.01666   0.00844  -0.0376   0.2980   0.9581
   3.250   0.5558   0.01848   0.00929  -0.0362   0.1843   0.9619
   3.500   0.5800   0.01955   0.01011  -0.0358   0.1557   0.9657
   3.750   0.6080   0.02057   0.01102  -0.0361   0.1375   0.9694
   4.250   0.6707   0.02284   0.01317  -0.0379   0.1140   0.9774
   4.500   0.7048   0.02436   0.01475  -0.0391   0.1059   0.9825
   4.750   0.7392   0.02596   0.01628  -0.0408   0.0974   0.9885
   5.000   0.7722   0.02771   0.01837  -0.0417   0.0922   1.0000
   5.250   0.8029   0.02929   0.02009  -0.0426   0.0861   1.0000
   5.500   0.8339   0.03258   0.02345  -0.0440   0.0827   1.0000
   5.750   0.8611   0.03491   0.02634  -0.0438   0.0813   1.0000
   6.000   0.8859   0.03760   0.02960  -0.0435   0.0793   1.0000
   6.250   0.9081   0.04081   0.03335  -0.0430   0.0780   1.0000
   6.500   0.9278   0.04498   0.03800  -0.0424   0.0794   1.0000
   6.750   0.9464   0.04978   0.04308  -0.0423   0.0815   1.0000
   7.000   0.8966   0.07201   0.06791  -0.0400   0.1799   1.0000
   7.250   0.9197   0.07584   0.07169  -0.0401   0.1741   1.0000
   7.500   0.8913   0.08115   0.07728  -0.0415   0.1649   1.0000
   7.750   0.9080   0.08479   0.08094  -0.0415   0.1593   1.0000
   8.000   0.9037   0.09045   0.08664  -0.0423   0.1550   1.0000
   8.250   0.8517   0.09781   0.09413  -0.0496   0.1509   1.0000
   8.500   0.8475   0.10359   0.09990  -0.0540   0.1459   1.0000
   8.750   0.8934   0.10527   0.10156  -0.0462   0.1405   1.0000
   9.000   0.8401   0.11564   0.11192  -0.0615   0.1384   1.0000
   9.250   0.8310   0.12150   0.11775  -0.0671   0.1322   1.0000
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