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NACA 23015 (naca23015-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 23015 (naca23015-il)
Reynolds number: 50,000
Max Cl/Cd: 20.95 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca23015-il-50000.txt
Download as CSV file: xf-naca23015-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 23015                                      
  
 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  -0.6857   0.08043   0.07310  -0.0302   0.9999   0.1960
  -9.750  -0.6731   0.07581   0.06848  -0.0301   0.9999   0.1935
  -9.500  -0.7549   0.06654   0.05860  -0.0296   0.9999   0.1797
  -9.250  -0.7533   0.06214   0.05400  -0.0286   0.9999   0.1784
  -9.000  -0.7509   0.05794   0.04956  -0.0274   0.9999   0.1773
  -8.750  -0.7479   0.05400   0.04531  -0.0259   0.9999   0.1764
  -8.500  -0.7431   0.05046   0.04137  -0.0243   0.9999   0.1766
  -8.250  -0.7341   0.04731   0.03780  -0.0227   0.9999   0.1776
  -8.000  -0.7216   0.04446   0.03452  -0.0211   0.9999   0.1784
  -7.750  -0.7060   0.04181   0.03153  -0.0197   0.9999   0.1802
  -7.500  -0.6854   0.03949   0.02924  -0.0187   0.9999   0.1841
  -7.250  -0.6656   0.03749   0.02707  -0.0174   0.9999   0.1878
  -7.000  -0.6457   0.03565   0.02496  -0.0161   0.9999   0.1916
  -6.750  -0.6259   0.03404   0.02302  -0.0146   0.9999   0.1964
  -6.500  -0.6028   0.03236   0.02148  -0.0137   0.9999   0.2024
  -6.250  -0.5803   0.03103   0.02004  -0.0125   0.9999   0.2094
  -6.000  -0.5572   0.02973   0.01882  -0.0114   0.9999   0.2181
  -5.750  -0.5343   0.02866   0.01771  -0.0101   0.9999   0.2289
  -5.500  -0.5117   0.02758   0.01682  -0.0089   0.9999   0.2429
  -5.250  -0.4907   0.02658   0.01598  -0.0075   0.9999   0.2635
  -5.000  -0.4725   0.02544   0.01525  -0.0058   0.9999   0.2962
  -4.750  -0.4612   0.02383   0.01465  -0.0031   0.9999   0.3918
  -4.500  -0.4619   0.02323   0.01533   0.0043   0.9999   0.6033
  -4.250  -0.4538   0.02392   0.01622   0.0105   0.9999   0.6909
  -4.000  -0.4397   0.02481   0.01716   0.0157   0.9999   0.7412
  -3.750  -0.4235   0.02562   0.01793   0.0201   0.9999   0.7819
  -3.500  -0.4024   0.02663   0.01888   0.0239   0.9999   0.8222
  -3.250  -0.2024   0.03245   0.02407   0.0083   0.9999   0.9085
  -3.000  -0.1513   0.03217   0.02362   0.0036   0.9999   0.9301
  -2.750  -0.1045   0.03179   0.02311  -0.0011   0.9999   0.9508
  -2.500  -0.0513   0.03137   0.02257  -0.0070   0.9999   0.9708
  -2.250   0.0089   0.03088   0.02197  -0.0145   0.9999   0.9906
  -2.000   0.0338   0.03057   0.02162  -0.0171   0.9999   1.0001
  -1.750   0.0151   0.03056   0.02163  -0.0133   0.9999   1.0001
  -1.500  -0.0074   0.03060   0.02169  -0.0092   0.9999   1.0001
  -1.250  -0.0313   0.03065   0.02175  -0.0051   0.9999   1.0001
  -1.000  -0.0544   0.03070   0.02181  -0.0010   0.9999   1.0001
  -0.750  -0.0295   0.03082   0.02186  -0.0039   0.9896   1.0001
  -0.500   0.0478   0.03083   0.02175  -0.0139   0.9651   1.0001
  -0.250   0.1278   0.03051   0.02136  -0.0235   0.9396   1.0001
   0.000   0.2030   0.02991   0.02071  -0.0316   0.9119   1.0001
   0.250   0.2783   0.02892   0.01969  -0.0389   0.8856   1.0001
   0.500   0.3275   0.02814   0.01888  -0.0416   0.8576   1.0001
   0.750   0.3635   0.02751   0.01821  -0.0420   0.8288   1.0001
   1.000   0.3899   0.02709   0.01772  -0.0407   0.8004   1.0001
   1.250   0.4120   0.02677   0.01732  -0.0386   0.7726   1.0001
   1.500   0.4314   0.02659   0.01704  -0.0362   0.7454   1.0001
   1.750   0.4491   0.02654   0.01690  -0.0337   0.7175   1.0001
   2.000   0.4686   0.02642   0.01664  -0.0311   0.6929   1.0001
   2.250   0.4850   0.02668   0.01681  -0.0288   0.6657   1.0001
   2.500   0.5032   0.02685   0.01686  -0.0265   0.6410   1.0001
   2.750   0.5225   0.02706   0.01693  -0.0243   0.6186   1.0001
   3.000   0.5422   0.02735   0.01707  -0.0223   0.5976   1.0001
   3.250   0.5609   0.02788   0.01750  -0.0205   0.5768   1.0001
   3.500   0.5803   0.02845   0.01798  -0.0189   0.5579   1.0001
   3.750   0.6006   0.02903   0.01846  -0.0174   0.5410   1.0001
   4.000   0.6192   0.02986   0.01927  -0.0160   0.5242   1.0001
   4.250   0.6375   0.03084   0.02027  -0.0148   0.5089   1.0001
   4.500   0.6566   0.03185   0.02128  -0.0136   0.4958   1.0001
   4.750   0.6799   0.03245   0.02174  -0.0124   0.4846   1.0001
   5.000   0.6956   0.03391   0.02336  -0.0113   0.4717   1.0001
   5.250   0.7150   0.03511   0.02457  -0.0103   0.4615   1.0001
   5.500   0.7338   0.03639   0.02590  -0.0094   0.4513   1.0001
   5.750   0.7497   0.03804   0.02766  -0.0084   0.4419   1.0001
   6.000   0.7680   0.03942   0.02910  -0.0074   0.4327   1.0001
   6.250   0.7827   0.04134   0.03112  -0.0065   0.4247   1.0001
   6.500   0.7918   0.04370   0.03365  -0.0054   0.4162   1.0001
   6.750   0.8167   0.04463   0.03454  -0.0046   0.4088   1.0001
   7.000   0.8023   0.04930   0.03955  -0.0032   0.4014   1.0001
   7.250   0.8258   0.05032   0.04057  -0.0023   0.3941   1.0001
   7.500   0.8076   0.05542   0.04588  -0.0009   0.3884   1.0001
   7.750   0.7236   0.06697   0.05757  -0.0004   0.3877   1.0001
   8.000   0.6194   0.08247   0.07299  -0.0053   0.3967   1.0001
   8.250   0.6399   0.08449   0.07506  -0.0046   0.3903   1.0001
   8.500   0.6261   0.09049   0.08108  -0.0063   0.3928   1.0001
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