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NACA 0008 (naca0008-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 0008 (naca0008-il)
Reynolds number: 50,000
Max Cl/Cd: 24.63 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca0008-il-50000-n5.txt
Download as CSV file: xf-naca0008-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0008                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.7159   0.09334   0.08645   0.0017   1.0000   0.0512
  -9.500  -0.7263   0.08537   0.07852  -0.0047   1.0000   0.0504
  -9.250  -0.7428   0.07833   0.07146  -0.0100   1.0000   0.0495
  -9.000  -0.7590   0.07249   0.06552  -0.0126   1.0000   0.0490
  -8.750  -0.7702   0.06691   0.05974  -0.0143   1.0000   0.0488
  -8.500  -0.7765   0.06165   0.05420  -0.0151   1.0000   0.0490
  -8.250  -0.7778   0.05665   0.04882  -0.0152   1.0000   0.0493
  -8.000  -0.7743   0.05197   0.04368  -0.0148   1.0000   0.0501
  -7.750  -0.7665   0.04777   0.03894  -0.0141   1.0000   0.0522
  -7.500  -0.7556   0.04383   0.03422  -0.0131   1.0000   0.0552
  -7.250  -0.7397   0.04054   0.03065  -0.0124   1.0000   0.0579
  -7.000  -0.7207   0.03779   0.02759  -0.0115   1.0000   0.0608
  -6.500  -0.6786   0.03305   0.02200  -0.0098   1.0000   0.0716
  -6.250  -0.6552   0.03106   0.01981  -0.0090   1.0000   0.0764
  -6.000  -0.6318   0.02939   0.01787  -0.0083   1.0000   0.0848
  -5.750  -0.6077   0.02780   0.01617  -0.0075   1.0000   0.0913
  -5.500  -0.5830   0.02627   0.01444  -0.0065   1.0000   0.0979
  -5.250  -0.5596   0.02494   0.01300  -0.0056   1.0000   0.1068
  -5.000  -0.5365   0.02373   0.01175  -0.0048   1.0000   0.1225
  -4.750  -0.5138   0.02245   0.01060  -0.0041   1.0000   0.1485
  -4.500  -0.4914   0.02105   0.00945  -0.0034   1.0000   0.1926
  -4.250  -0.4710   0.01971   0.00860  -0.0026   1.0000   0.2797
  -4.000  -0.4535   0.01856   0.00809  -0.0009   1.0000   0.4032
  -3.750  -0.4364   0.01772   0.00782   0.0017   1.0000   0.5239
  -3.500  -0.4176   0.01718   0.00764   0.0045   1.0000   0.6200
  -3.250  -0.3967   0.01685   0.00752   0.0073   1.0000   0.7031
  -3.000  -0.3717   0.01671   0.00743   0.0095   1.0000   0.7747
  -2.750  -0.3403   0.01668   0.00732   0.0103   1.0000   0.8347
  -2.500  -0.2999   0.01670   0.00713   0.0090   1.0000   0.8838
  -2.250  -0.2537   0.01669   0.00688   0.0060   1.0000   0.9244
  -2.000  -0.1993   0.01660   0.00653   0.0008   1.0000   0.9563
  -1.750  -0.1431   0.01640   0.00607  -0.0052   1.0000   0.9836
  -1.500  -0.0988   0.01611   0.00563  -0.0093   1.0000   1.0000
  -1.250  -0.0818   0.01588   0.00532  -0.0081   1.0000   1.0000
  -1.000  -0.0649   0.01571   0.00508  -0.0067   1.0000   1.0000
  -0.750  -0.0483   0.01557   0.00491  -0.0052   1.0000   1.0000
  -0.500  -0.0319   0.01548   0.00479  -0.0035   1.0000   1.0000
  -0.250  -0.0158   0.01543   0.00471  -0.0018   1.0000   1.0000
   0.000   0.0000   0.01541   0.00469   0.0000   1.0000   1.0000
   0.250   0.0158   0.01543   0.00471   0.0018   1.0000   1.0000
   0.500   0.0319   0.01548   0.00479   0.0035   1.0000   1.0000
   0.750   0.0483   0.01557   0.00491   0.0052   1.0000   1.0000
   1.000   0.0649   0.01570   0.00508   0.0067   1.0000   1.0000
   1.250   0.0818   0.01588   0.00532   0.0081   1.0000   1.0000
   1.500   0.0988   0.01611   0.00563   0.0093   1.0000   1.0000
   1.750   0.1431   0.01640   0.00607   0.0052   0.9836   1.0000
   2.000   0.1993   0.01660   0.00653  -0.0008   0.9563   1.0000
   2.250   0.2537   0.01669   0.00688  -0.0060   0.9244   1.0000
   2.500   0.2999   0.01670   0.00713  -0.0090   0.8839   1.0000
   2.750   0.3403   0.01668   0.00732  -0.0103   0.8348   1.0000
   3.000   0.3717   0.01671   0.00743  -0.0095   0.7747   1.0000
   3.250   0.3967   0.01685   0.00752  -0.0073   0.7031   1.0000
   3.500   0.4176   0.01718   0.00764  -0.0045   0.6200   1.0000
   3.750   0.4364   0.01772   0.00782  -0.0017   0.5239   1.0000
   4.000   0.4535   0.01856   0.00809   0.0009   0.4033   1.0000
   4.250   0.4710   0.01971   0.00860   0.0026   0.2797   1.0000
   4.500   0.4915   0.02105   0.00945   0.0034   0.1926   1.0000
   4.750   0.5139   0.02245   0.01060   0.0041   0.1485   1.0000
   5.000   0.5365   0.02373   0.01175   0.0048   0.1225   1.0000
   5.250   0.5597   0.02494   0.01300   0.0056   0.1068   1.0000
   5.500   0.5830   0.02627   0.01444   0.0065   0.0979   1.0000
   5.750   0.6077   0.02780   0.01617   0.0075   0.0913   1.0000
   6.000   0.6318   0.02939   0.01787   0.0083   0.0848   1.0000
   6.250   0.6553   0.03106   0.01981   0.0090   0.0764   1.0000
   6.500   0.6786   0.03305   0.02200   0.0098   0.0716   1.0000
   7.000   0.7207   0.03779   0.02759   0.0115   0.0608   1.0000
   7.250   0.7397   0.04054   0.03065   0.0124   0.0579   1.0000
   7.500   0.7556   0.04383   0.03422   0.0131   0.0552   1.0000
   7.750   0.7665   0.04777   0.03894   0.0141   0.0522   1.0000
   8.000   0.7743   0.05197   0.04368   0.0148   0.0501   1.0000
   8.250   0.7778   0.05666   0.04882   0.0152   0.0493   1.0000
   8.500   0.7765   0.06166   0.05420   0.0151   0.0490   1.0000
   8.750   0.7703   0.06692   0.05975   0.0143   0.0488   1.0000
   9.000   0.7591   0.07250   0.06553   0.0126   0.0490   1.0000
   9.250   0.7430   0.07836   0.07149   0.0100   0.0495   1.0000
   9.500   0.7266   0.08542   0.07856   0.0046   0.0504   1.0000
   9.750   0.7163   0.09339   0.08650  -0.0018   0.0512   1.0000
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