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NASA SC(2)-0714 AIRFOIL (sc20714-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA SC(2)-0714 AIRFOIL (sc20714-il)
Reynolds number: 50,000
Max Cl/Cd: 25.43 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sc20714-il-50000.txt
Download as CSV file: xf-sc20714-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA SC(2)-0714 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6806   0.08160   0.07361  -0.0435   1.0000   0.1795
  -9.750  -0.7107   0.07606   0.06804  -0.0434   1.0000   0.1755
  -9.500  -0.7340   0.07057   0.06248  -0.0438   1.0000   0.1730
  -9.250  -0.7581   0.06467   0.05636  -0.0450   1.0000   0.1704
  -9.000  -0.7691   0.05926   0.05059  -0.0464   1.0000   0.1693
  -8.750  -0.7660   0.05481   0.04578  -0.0476   1.0000   0.1703
  -8.500  -0.7550   0.05076   0.04133  -0.0490   1.0000   0.1720
  -8.250  -0.7379   0.04701   0.03711  -0.0506   1.0000   0.1739
  -8.000  -0.7151   0.04368   0.03310  -0.0530   1.0000   0.1778
  -7.750  -0.6947   0.04137   0.03089  -0.0524   1.0000   0.1829
  -7.500  -0.6711   0.03924   0.02854  -0.0526   1.0000   0.1885
  -7.250  -0.6447   0.03706   0.02598  -0.0536   1.0000   0.1958
  -7.000  -0.6226   0.03551   0.02451  -0.0526   1.0000   0.2040
  -6.750  -0.5977   0.03389   0.02270  -0.0524   1.0000   0.2132
  -6.500  -0.5746   0.03265   0.02144  -0.0514   1.0000   0.2256
  -6.250  -0.5543   0.03155   0.02051  -0.0494   1.0000   0.2387
  -6.000  -0.5336   0.03055   0.01963  -0.0475   1.0000   0.2567
  -5.750  -0.5141   0.02956   0.01889  -0.0453   1.0000   0.2803
  -5.500  -0.4932   0.02837   0.01817  -0.0439   1.0000   0.3192
  -5.250  -0.4743   0.02700   0.01802  -0.0418   1.0000   0.4103
  -5.000  -0.4843   0.02934   0.02141  -0.0281   1.0000   0.4976
  -4.750  -0.4790   0.03331   0.02534  -0.0177   1.0000   0.5865
  -4.500  -0.4694   0.03617   0.02801  -0.0100   1.0000   0.6353
  -4.250  -0.4641   0.03831   0.03008  -0.0008   1.0000   0.6619
  -4.000  -0.4543   0.03970   0.03135   0.0058   1.0000   0.6931
  -3.750  -0.4451   0.04069   0.03225   0.0130   1.0000   0.7208
  -3.500  -0.4341   0.04105   0.03250   0.0182   1.0000   0.7506
  -3.250  -0.4222   0.04105   0.03240   0.0233   1.0000   0.7797
  -3.000  -0.4081   0.04065   0.03191   0.0281   1.0000   0.8098
  -2.750  -0.3906   0.03995   0.03111   0.0317   1.0000   0.8425
  -2.500  -0.3689   0.03905   0.03008   0.0340   1.0000   0.8769
  -2.250  -0.2527   0.03784   0.02854   0.0204   1.0000   0.9311
  -2.000  -0.1761   0.03674   0.02725   0.0104   1.0000   0.9574
  -1.750  -0.1339   0.03595   0.02636   0.0060   1.0000   0.9698
  -1.500  -0.0936   0.03528   0.02564   0.0017   1.0000   0.9791
  -1.250  -0.0597   0.03479   0.02510  -0.0014   1.0000   0.9865
  -1.000  -0.0320   0.03446   0.02475  -0.0034   1.0000   0.9930
  -0.750  -0.0003   0.03418   0.02447  -0.0061   1.0000   0.9989
  -0.500   0.0090   0.03399   0.02432  -0.0046   1.0000   1.0000
  -0.250   0.0135   0.03383   0.02419  -0.0023   1.0000   1.0000
   0.000   0.0176   0.03370   0.02410   0.0001   1.0000   1.0000
   0.250   0.0215   0.03360   0.02404   0.0025   1.0000   1.0000
   0.500   0.0250   0.03352   0.02401   0.0049   1.0000   1.0000
   0.750   0.0281   0.03347   0.02402   0.0073   1.0000   1.0000
   1.000   0.0308   0.03345   0.02405   0.0097   1.0000   1.0000
   1.250   0.0331   0.03345   0.02412   0.0121   1.0000   1.0000
   1.500   0.0350   0.03348   0.02423   0.0144   1.0000   1.0000
   1.750   0.0365   0.03355   0.02436   0.0167   1.0000   1.0000
   2.000   0.0378   0.03365   0.02455   0.0189   1.0000   1.0000
   2.250   0.0390   0.03380   0.02479   0.0209   1.0000   1.0000
   2.500   0.0404   0.03402   0.02510   0.0227   1.0000   1.0000
   2.750   0.0589   0.03477   0.02597   0.0213   0.9951   1.0000
   3.000   0.1673   0.03722   0.02862   0.0049   0.9534   1.0000
   3.250   0.2741   0.03759   0.02926  -0.0085   0.9041   1.0000
   3.500   0.3745   0.03610   0.02810  -0.0183   0.8559   1.0000
   3.750   0.4634   0.03316   0.02555  -0.0242   0.8073   1.0000
   4.000   0.6513   0.02561   0.01638  -0.0338   0.4163   1.0000
   4.250   0.6671   0.02747   0.01729  -0.0316   0.3285   1.0000
   4.500   0.7061   0.02904   0.01832  -0.0337   0.2790   1.0000
   4.750   0.7421   0.03034   0.01950  -0.0352   0.2509   1.0000
   5.000   0.7734   0.03167   0.02074  -0.0359   0.2325   1.0000
   5.250   0.8008   0.03312   0.02222  -0.0360   0.2201   1.0000
   5.500   0.8264   0.03469   0.02364  -0.0360   0.2093   1.0000
   5.750   0.8405   0.03598   0.02528  -0.0337   0.2019   1.0000
   6.000   0.8604   0.03756   0.02686  -0.0327   0.1950   1.0000
   6.250   0.8728   0.03929   0.02884  -0.0304   0.1896   1.0000
   6.500   0.8825   0.04094   0.03083  -0.0277   0.1847   1.0000
   6.750   0.8988   0.04286   0.03290  -0.0263   0.1800   1.0000
   7.000   0.9180   0.04542   0.03551  -0.0258   0.1752   1.0000
   7.250   0.9254   0.04787   0.03851  -0.0234   0.1725   1.0000
   7.500   0.9345   0.05085   0.04197  -0.0216   0.1705   1.0000
   7.750   0.9416   0.05408   0.04565  -0.0200   0.1685   1.0000
   8.000   0.9468   0.05756   0.04952  -0.0186   0.1668   1.0000
   8.250   0.9464   0.06156   0.05393  -0.0170   0.1664   1.0000
   8.500   0.9345   0.06642   0.05925  -0.0151   0.1685   1.0000
   8.750   0.9209   0.07162   0.06478  -0.0139   0.1713   1.0000
   9.000   0.9118   0.07687   0.07025  -0.0134   0.1737   1.0000
   9.250   0.9169   0.08231   0.07579  -0.0141   0.1756   1.0000
   9.500   0.7502   0.10712   0.10115  -0.0288   0.2261   1.0000
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