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NACA 16-012 (naca16012-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 16-012 (naca16012-il)
Reynolds number: 100,000
Max Cl/Cd: 28.15 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16012-il-100000.txt
Download as CSV file: xf-naca16012-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16-012                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.5855   0.11045   0.10530  -0.0337   1.0000   0.1256
 -10.750  -0.6067   0.10569   0.10062  -0.0366   1.0000   0.1304
 -10.500  -0.6510   0.09948   0.09453  -0.0421   1.0000   0.1317
 -10.250  -0.6324   0.09605   0.09111  -0.0385   1.0000   0.1362
 -10.000  -0.6385   0.09245   0.08755  -0.0379   1.0000   0.1405
  -9.750  -0.6695   0.08810   0.08323  -0.0380   1.0000   0.1426
  -9.500  -0.7059   0.08501   0.08015  -0.0351   1.0000   0.1438
  -9.250  -0.7439   0.08305   0.07820  -0.0293   1.0000   0.1445
  -9.000  -0.7883   0.08164   0.07670  -0.0223   1.0000   0.1457
  -8.750  -0.8373   0.08137   0.07620  -0.0139   1.0000   0.1467
  -8.500  -0.7923   0.07442   0.06953  -0.0164   1.0000   0.1548
  -8.250  -0.8403   0.07423   0.06903  -0.0078   1.0000   0.1606
  -7.000  -0.8670   0.04776   0.03997   0.0179   1.0000   0.0760
  -6.750  -0.8573   0.04375   0.03571   0.0207   1.0000   0.0744
  -6.500  -0.8456   0.04020   0.03180   0.0238   1.0000   0.0717
  -6.250  -0.8314   0.03703   0.02813   0.0269   1.0000   0.0698
  -6.000  -0.8139   0.03454   0.02520   0.0294   1.0000   0.0701
  -5.750  -0.7951   0.03282   0.02309   0.0314   1.0000   0.0736
  -5.500  -0.7729   0.03106   0.02091   0.0331   1.0000   0.0756
  -5.250  -0.7456   0.02890   0.01849   0.0336   1.0000   0.0782
  -5.000  -0.7216   0.02752   0.01711   0.0341   1.0000   0.0849
  -4.750  -0.6963   0.02639   0.01575   0.0350   1.0000   0.0898
  -4.500  -0.6714   0.02492   0.01443   0.0354   1.0000   0.0974
  -4.250  -0.6500   0.02391   0.01338   0.0367   1.0000   0.1061
  -4.000  -0.6320   0.02292   0.01245   0.0386   1.0000   0.1160
  -3.750  -0.6165   0.02199   0.01161   0.0409   1.0000   0.1332
  -3.500  -0.5583   0.01984   0.01342   0.0395   1.0000   0.8530
  -3.250  -0.1899   0.02718   0.01937  -0.0189   1.0000   0.9851
  -3.000  -0.1166   0.02642   0.01834  -0.0296   1.0000   1.0000
  -2.750  -0.1072   0.02610   0.01795  -0.0273   1.0000   1.0000
  -2.500  -0.0978   0.02582   0.01761  -0.0249   1.0000   1.0000
  -2.250  -0.0882   0.02558   0.01730  -0.0225   1.0000   1.0000
  -2.000  -0.0786   0.02536   0.01704  -0.0201   1.0000   1.0000
  -1.750  -0.0689   0.02518   0.01680  -0.0176   1.0000   1.0000
  -1.500  -0.0592   0.02503   0.01661  -0.0151   1.0000   1.0000
  -1.250  -0.0494   0.02490   0.01644  -0.0126   1.0000   1.0000
  -1.000  -0.0395   0.02479   0.01631  -0.0101   1.0000   1.0000
  -0.750  -0.0297   0.02471   0.01621  -0.0076   1.0000   1.0000
  -0.500  -0.0198   0.02465   0.01614  -0.0051   1.0000   1.0000
  -0.250  -0.0099   0.02462   0.01609  -0.0025   1.0000   1.0000
   0.000   0.0000   0.02460   0.01608   0.0000   1.0000   1.0000
   0.250   0.0099   0.02462   0.01609   0.0025   1.0000   1.0000
   0.500   0.0198   0.02465   0.01613   0.0051   1.0000   1.0000
   0.750   0.0297   0.02470   0.01621   0.0076   1.0000   1.0000
   1.000   0.0395   0.02478   0.01631   0.0101   1.0000   1.0000
   1.250   0.0494   0.02489   0.01643   0.0126   1.0000   1.0000
   1.500   0.0592   0.02501   0.01660   0.0151   1.0000   1.0000
   1.750   0.0689   0.02517   0.01679   0.0176   1.0000   1.0000
   2.000   0.0786   0.02535   0.01703   0.0201   1.0000   1.0000
   2.250   0.0882   0.02556   0.01729   0.0225   1.0000   1.0000
   2.500   0.0978   0.02580   0.01759   0.0249   1.0000   1.0000
   2.750   0.1072   0.02608   0.01793   0.0273   1.0000   1.0000
   3.000   0.1166   0.02640   0.01831   0.0296   1.0000   1.0000
   3.250   0.1895   0.02716   0.01934   0.0190   0.9852   1.0000
   3.500   0.5583   0.01983   0.01341  -0.0395   0.8530   1.0000
   3.750   0.6164   0.02198   0.01160  -0.0409   0.1333   1.0000
   4.000   0.6318   0.02291   0.01244  -0.0386   0.1160   1.0000
   4.250   0.6498   0.02390   0.01337  -0.0367   0.1061   1.0000
   4.500   0.6712   0.02491   0.01442  -0.0354   0.0974   1.0000
   4.750   0.6961   0.02638   0.01574  -0.0350   0.0898   1.0000
   5.000   0.7214   0.02751   0.01710  -0.0341   0.0849   1.0000
   5.250   0.7454   0.02888   0.01847  -0.0336   0.0783   1.0000
   5.500   0.7727   0.03105   0.02089  -0.0331   0.0756   1.0000
   5.750   0.7950   0.03281   0.02308  -0.0314   0.0736   1.0000
   6.000   0.8137   0.03453   0.02519  -0.0293   0.0701   1.0000
   6.250   0.8313   0.03701   0.02811  -0.0269   0.0698   1.0000
   6.500   0.8455   0.04018   0.03177  -0.0238   0.0717   1.0000
   6.750   0.8572   0.04373   0.03569  -0.0207   0.0744   1.0000
   7.000   0.8669   0.04774   0.03995  -0.0179   0.0760   1.0000
   8.250   0.8410   0.07427   0.06906   0.0078   0.1606   1.0000
   8.750   0.8372   0.08135   0.07617   0.0139   0.1467   1.0000
   9.000   0.7881   0.08161   0.07667   0.0223   0.1457   1.0000
   9.250   0.7437   0.08303   0.07817   0.0293   0.1444   1.0000
   9.500   0.7058   0.08498   0.08013   0.0351   0.1437   1.0000
   9.750   0.6695   0.08809   0.08322   0.0380   0.1425   1.0000
  10.000   0.6389   0.09242   0.08751   0.0379   0.1404   1.0000
  10.250   0.5031   0.09616   0.09147   0.0338   0.1571   1.0000
  10.500   0.4975   0.09967   0.09496   0.0344   0.1506   1.0000
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