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NACA 0012 AIRFOILS (n0012-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0012 AIRFOILS (n0012-il)
Reynolds number: 50,000
Max Cl/Cd: 25.67 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n0012-il-50000.txt
Download as CSV file: xf-n0012-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0012 AIRFOILS                              
  
 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  -0.8195   0.07577   0.06813  -0.0225   1.0000   0.1547
  -9.750  -0.8355   0.07025   0.06249  -0.0224   1.0000   0.1540
  -9.500  -0.8519   0.06488   0.05690  -0.0217   1.0000   0.1535
  -9.250  -0.8654   0.05977   0.05144  -0.0204   1.0000   0.1535
  -9.000  -0.8743   0.05500   0.04622  -0.0186   1.0000   0.1540
  -8.750  -0.8723   0.05095   0.04184  -0.0169   1.0000   0.1566
  -8.500  -0.8537   0.04867   0.03960  -0.0158   1.0000   0.1632
  -8.250  -0.8499   0.04511   0.03550  -0.0138   1.0000   0.1676
  -8.000  -0.8371   0.04218   0.03233  -0.0122   1.0000   0.1739
  -7.750  -0.8225   0.03983   0.02970  -0.0105   1.0000   0.1828
  -7.500  -0.8044   0.03756   0.02736  -0.0092   1.0000   0.1924
  -7.250  -0.7876   0.03540   0.02498  -0.0076   1.0000   0.2045
  -7.000  -0.7691   0.03350   0.02292  -0.0061   1.0000   0.2190
  -6.750  -0.7491   0.03186   0.02130  -0.0047   1.0000   0.2365
  -6.500  -0.7296   0.03038   0.01986  -0.0031   1.0000   0.2578
  -6.250  -0.7105   0.02892   0.01844  -0.0014   1.0000   0.2830
  -6.000  -0.6919   0.02765   0.01725   0.0005   1.0000   0.3133
  -5.750  -0.6730   0.02658   0.01641   0.0025   1.0000   0.3480
  -5.500  -0.6548   0.02564   0.01570   0.0047   1.0000   0.3875
  -5.250  -0.6369   0.02482   0.01509   0.0072   1.0000   0.4301
  -5.000  -0.6195   0.02413   0.01455   0.0098   1.0000   0.4753
  -4.750  -0.6002   0.02367   0.01438   0.0126   1.0000   0.5186
  -4.500  -0.5816   0.02327   0.01415   0.0155   1.0000   0.5624
  -4.250  -0.5628   0.02300   0.01400   0.0186   1.0000   0.6051
  -4.000  -0.5431   0.02285   0.01398   0.0218   1.0000   0.6455
  -3.750  -0.5227   0.02282   0.01405   0.0250   1.0000   0.6846
  -3.500  -0.5030   0.02279   0.01404   0.0283   1.0000   0.7237
  -3.250  -0.4793   0.02298   0.01425   0.0313   1.0000   0.7602
  -3.000  -0.4493   0.02334   0.01458   0.0333   1.0000   0.7958
  -2.750  -0.4085   0.02382   0.01495   0.0333   1.0000   0.8311
  -2.500  -0.3500   0.02437   0.01531   0.0298   1.0000   0.8657
  -2.250  -0.2694   0.02476   0.01543   0.0214   1.0000   0.8987
  -2.000  -0.1829   0.02464   0.01505   0.0107   1.0000   0.9296
  -1.750  -0.1037   0.02407   0.01431   0.0003   1.0000   0.9591
  -1.500  -0.0223   0.02312   0.01322  -0.0113   1.0000   0.9863
  -1.250   0.0255   0.02220   0.01226  -0.0174   1.0000   1.0000
  -1.000   0.0285   0.02168   0.01177  -0.0155   1.0000   1.0000
  -0.750   0.0275   0.02128   0.01142  -0.0127   1.0000   1.0000
  -0.500   0.0225   0.02100   0.01117  -0.0092   1.0000   1.0000
  -0.250   0.0128   0.02083   0.01103  -0.0049   1.0000   1.0000
   0.000   0.0000   0.02078   0.01099   0.0000   1.0000   1.0000
   0.250  -0.0128   0.02083   0.01103   0.0049   1.0000   1.0000
   0.500  -0.0225   0.02099   0.01117   0.0092   1.0000   1.0000
   0.750  -0.0275   0.02128   0.01141   0.0127   1.0000   1.0000
   1.000  -0.0285   0.02168   0.01177   0.0155   1.0000   1.0000
   1.250  -0.0255   0.02219   0.01225   0.0174   1.0000   1.0000
   1.500   0.0223   0.02311   0.01322   0.0113   0.9864   1.0000
   1.750   0.1036   0.02406   0.01430  -0.0003   0.9592   1.0000
   2.000   0.1829   0.02463   0.01505  -0.0106   0.9296   1.0000
   2.250   0.2694   0.02476   0.01543  -0.0214   0.8987   1.0000
   2.500   0.3499   0.02436   0.01530  -0.0298   0.8657   1.0000
   2.750   0.4085   0.02382   0.01494  -0.0333   0.8311   1.0000
   3.000   0.4492   0.02334   0.01458  -0.0333   0.7958   1.0000
   3.250   0.4792   0.02298   0.01425  -0.0313   0.7602   1.0000
   3.500   0.5030   0.02279   0.01404  -0.0283   0.7237   1.0000
   3.750   0.5227   0.02282   0.01405  -0.0250   0.6847   1.0000
   4.000   0.5431   0.02284   0.01398  -0.0218   0.6456   1.0000
   4.250   0.5627   0.02300   0.01400  -0.0186   0.6051   1.0000
   4.500   0.5816   0.02327   0.01415  -0.0155   0.5624   1.0000
   4.750   0.6001   0.02367   0.01438  -0.0126   0.5187   1.0000
   5.000   0.6194   0.02413   0.01455  -0.0098   0.4754   1.0000
   5.250   0.6368   0.02482   0.01509  -0.0072   0.4301   1.0000
   5.500   0.6547   0.02564   0.01570  -0.0047   0.3875   1.0000
   5.750   0.6730   0.02658   0.01641  -0.0025   0.3481   1.0000
   6.000   0.6919   0.02765   0.01724  -0.0005   0.3133   1.0000
   6.250   0.7105   0.02892   0.01844   0.0014   0.2830   1.0000
   6.500   0.7295   0.03038   0.01986   0.0031   0.2578   1.0000
   6.750   0.7491   0.03186   0.02130   0.0047   0.2365   1.0000
   7.000   0.7691   0.03350   0.02292   0.0061   0.2190   1.0000
   7.250   0.7876   0.03540   0.02498   0.0076   0.2045   1.0000
   7.500   0.8044   0.03756   0.02736   0.0092   0.1924   1.0000
   7.750   0.8226   0.03983   0.02970   0.0105   0.1828   1.0000
   8.000   0.8372   0.04218   0.03233   0.0122   0.1739   1.0000
   8.250   0.8499   0.04511   0.03550   0.0138   0.1676   1.0000
   8.500   0.8538   0.04867   0.03960   0.0158   0.1632   1.0000
   8.750   0.8727   0.05094   0.04181   0.0168   0.1565   1.0000
   9.000   0.8743   0.05501   0.04623   0.0185   0.1540   1.0000
   9.250   0.8655   0.05978   0.05145   0.0203   0.1535   1.0000
   9.500   0.8520   0.06490   0.05691   0.0216   0.1535   1.0000
   9.750   0.8357   0.07027   0.06251   0.0223   0.1540   1.0000
  10.000   0.8198   0.07580   0.06817   0.0224   0.1547   1.0000
  10.250   0.6499   0.10541   0.09769  -0.0021   0.2047   1.0000
  10.500   0.6743   0.10839   0.10074   0.0003   0.2006   1.0000
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