Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 23112 (naca23112-jf) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 23112 (naca23112-jf)
Reynolds number: 50,000
Max Cl/Cd: 27.32 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca23112-jf-50000-n5.txt
Download as CSV file: xf-naca23112-jf-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 23112                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.4493   0.13046   0.12352  -0.0066   1.0000   0.1374
 -11.500  -0.4474   0.12617   0.11927  -0.0074   1.0000   0.1379
 -11.250  -0.5767   0.12035   0.11294  -0.0042   1.0000   0.0772
 -11.000  -0.5568   0.11742   0.10995  -0.0014   1.0000   0.0736
 -10.500  -0.5817   0.10164   0.09424  -0.0139   1.0000   0.0646
 -10.250  -0.5832   0.09643   0.08905  -0.0166   1.0000   0.0642
 -10.000  -0.5897   0.09098   0.08362  -0.0198   1.0000   0.0637
  -9.750  -0.6004   0.08577   0.07839  -0.0226   1.0000   0.0633
  -9.500  -0.6141   0.08109   0.07368  -0.0241   1.0000   0.0629
  -9.250  -0.6285   0.07695   0.06946  -0.0242   1.0000   0.0626
  -9.000  -0.6393   0.07281   0.06519  -0.0241   1.0000   0.0625
  -8.750  -0.6466   0.06885   0.06104  -0.0234   1.0000   0.0625
  -8.500  -0.6504   0.06503   0.05701  -0.0224   1.0000   0.0625
  -8.250  -0.6507   0.06136   0.05307  -0.0212   1.0000   0.0624
  -8.000  -0.6479   0.05783   0.04926  -0.0197   1.0000   0.0623
  -7.750  -0.6426   0.05449   0.04560  -0.0181   1.0000   0.0623
  -7.500  -0.6349   0.05135   0.04211  -0.0163   1.0000   0.0625
  -7.250  -0.6259   0.04848   0.03881  -0.0143   1.0000   0.0635
  -7.000  -0.6157   0.04593   0.03570  -0.0120   1.0000   0.0649
  -6.750  -0.5997   0.04341   0.03311  -0.0108   1.0000   0.0664
  -6.500  -0.5827   0.04122   0.03075  -0.0094   1.0000   0.0676
  -6.250  -0.5648   0.03910   0.02839  -0.0079   1.0000   0.0686
  -6.000  -0.5461   0.03725   0.02631  -0.0064   1.0000   0.0706
  -5.750  -0.5267   0.03558   0.02437  -0.0049   1.0000   0.0734
  -5.500  -0.5062   0.03400   0.02243  -0.0033   1.0000   0.0759
  -5.250  -0.4838   0.03233   0.02057  -0.0021   1.0000   0.0776
  -5.000  -0.4613   0.03086   0.01914  -0.0012   1.0000   0.0807
  -4.750  -0.4386   0.02968   0.01792  -0.0001   1.0000   0.0849
  -4.500  -0.4146   0.02859   0.01668   0.0010   1.0000   0.0886
  -4.250  -0.3906   0.02745   0.01556   0.0020   1.0000   0.0926
  -4.000  -0.3688   0.02654   0.01471   0.0031   1.0000   0.0983
  -3.750  -0.3474   0.02579   0.01386   0.0045   1.0000   0.1036
  -3.500  -0.3283   0.02496   0.01308   0.0059   1.0000   0.1099
  -3.250  -0.3103   0.02431   0.01241   0.0075   1.0000   0.1180
  -3.000  -0.2932   0.02363   0.01175   0.0090   1.0000   0.1265
  -2.750  -0.2759   0.02303   0.01116   0.0104   1.0000   0.1394
  -2.500  -0.2535   0.02222   0.01053   0.0106   0.9966   0.1623
  -2.250  -0.2286   0.01929   0.01030   0.0106   0.9838   0.6073
  -2.000  -0.1757   0.01961   0.01126   0.0102   0.9735   0.8429
  -1.750  -0.0857   0.02055   0.01195   0.0017   0.9679   0.9332
  -1.500   0.0322   0.02074   0.01178  -0.0140   0.9608   0.9815
  -1.250   0.0971   0.02035   0.01118  -0.0215   0.9301   0.9898
  -1.000   0.1563   0.01994   0.01058  -0.0277   0.8944   0.9987
  -0.750   0.1918   0.01966   0.01012  -0.0292   0.8547   1.0000
  -0.500   0.2176   0.01946   0.00975  -0.0287   0.8151   1.0000
  -0.250   0.2403   0.01933   0.00942  -0.0274   0.7773   1.0000
   0.000   0.2613   0.01927   0.00914  -0.0258   0.7413   1.0000
   0.250   0.2818   0.01926   0.00892  -0.0241   0.7066   1.0000
   0.750   0.3222   0.01941   0.00864  -0.0209   0.6403   1.0000
   1.000   0.3426   0.01954   0.00858  -0.0194   0.6091   1.0000
   1.250   0.3632   0.01972   0.00856  -0.0179   0.5793   1.0000
   1.500   0.3839   0.01994   0.00858  -0.0165   0.5516   1.0000
   1.750   0.4047   0.02018   0.00864  -0.0152   0.5257   1.0000
   2.000   0.4258   0.02047   0.00876  -0.0140   0.5019   1.0000
   2.250   0.4469   0.02077   0.00893  -0.0128   0.4795   1.0000
   2.500   0.4683   0.02111   0.00912  -0.0116   0.4597   1.0000
   2.750   0.4905   0.02148   0.00935  -0.0107   0.4418   1.0000
   3.000   0.5128   0.02187   0.00965  -0.0097   0.4253   1.0000
   3.250   0.5351   0.02228   0.00998  -0.0088   0.4100   1.0000
   3.500   0.5574   0.02272   0.01036  -0.0079   0.3963   1.0000
   3.750   0.5798   0.02318   0.01077  -0.0070   0.3839   1.0000
   4.000   0.6022   0.02366   0.01115  -0.0061   0.3730   1.0000
   4.250   0.6244   0.02417   0.01167  -0.0052   0.3618   1.0000
   4.500   0.6468   0.02472   0.01223  -0.0043   0.3520   1.0000
   4.750   0.6691   0.02525   0.01271  -0.0034   0.3430   1.0000
   5.000   0.6910   0.02588   0.01340  -0.0025   0.3336   1.0000
   5.250   0.7135   0.02646   0.01391  -0.0016   0.3260   1.0000
   5.500   0.7348   0.02716   0.01477  -0.0007   0.3173   1.0000
   5.750   0.7572   0.02778   0.01534   0.0001   0.3102   1.0000
   6.000   0.7781   0.02858   0.01628   0.0010   0.3023   1.0000
   6.250   0.8000   0.02928   0.01702   0.0019   0.2955   1.0000
   6.500   0.8208   0.03013   0.01798   0.0028   0.2884   1.0000
   6.750   0.8417   0.03095   0.01890   0.0037   0.2813   1.0000
   7.000   0.8636   0.03176   0.01972   0.0045   0.2754   1.0000
   7.250   0.8820   0.03283   0.02105   0.0056   0.2681   1.0000
   7.500   0.9042   0.03358   0.02179   0.0063   0.2623   1.0000
   7.750   0.9213   0.03483   0.02329   0.0074   0.2555   1.0000
   8.000   0.9406   0.03584   0.02445   0.0084   0.2492   1.0000
   8.250   0.9612   0.03678   0.02543   0.0092   0.2436   1.0000
   8.500   0.9746   0.03830   0.02727   0.0105   0.2366   1.0000
   8.750   0.9958   0.03914   0.02814   0.0113   0.2311   1.0000
   9.000   1.0076   0.04080   0.03011   0.0126   0.2248   1.0000
   9.250   1.0218   0.04216   0.03166   0.0138   0.2185   1.0000
   9.500   1.0437   0.04292   0.03239   0.0145   0.2134   1.0000
   9.750   1.0435   0.04546   0.03538   0.0165   0.2067   1.0000
  10.000   1.0610   0.04637   0.03638   0.0175   0.2010   1.0000
  10.250   1.0660   0.04844   0.03868   0.0190   0.1956   1.0000
  10.500   1.0621   0.05108   0.04162   0.0208   0.1899   1.0000
  10.750   1.0896   0.05096   0.04143   0.0215   0.1844   1.0000
  11.000   1.0620   0.05552   0.04638   0.0237   0.1802   1.0000
  11.250   1.0221   0.06098   0.05208   0.0254   0.1772   1.0000
  11.500   0.9132   0.07824   0.06943   0.0172   0.1750   1.0000
  11.750   1.0200   0.06586   0.05709   0.0261   0.1685   1.0000
  12.000   0.8240   0.10265   0.09378   0.0037   0.1627   1.0000
  12.250   0.8305   0.10511   0.09630   0.0035   0.1581   1.0000
<< Back to NACA 23112 (naca23112-jf)

Polar data table (+)

Polar graphs


<< Back to NACA 23112 (naca23112-jf)