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NACA 23112 (naca23112-jf) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 23112 (naca23112-jf)
Reynolds number: 50,000
Max Cl/Cd: 22.35 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca23112-jf-50000.txt
Download as CSV file: xf-naca23112-jf-50000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4638   0.10226   0.09542   0.0252   1.0000   0.4460
  -8.250  -0.4369   0.09805   0.09118   0.0261   1.0000   0.4697
  -8.000  -0.4516   0.09702   0.09026   0.0282   1.0000   0.4942
  -7.500  -0.4117   0.08827   0.08152   0.0277   1.0000   0.5189
  -6.750  -0.6083   0.05479   0.04716  -0.0121   1.0000   0.2005
  -6.500  -0.5966   0.05015   0.04210  -0.0113   1.0000   0.1819
  -6.250  -0.5864   0.04666   0.03790  -0.0095   1.0000   0.1683
  -6.000  -0.5700   0.04331   0.03434  -0.0080   1.0000   0.1630
  -5.750  -0.5557   0.04132   0.03139  -0.0053   1.0000   0.1537
  -5.500  -0.5357   0.03840   0.02831  -0.0040   1.0000   0.1518
  -5.250  -0.5161   0.03626   0.02585  -0.0023   1.0000   0.1522
  -5.000  -0.4954   0.03438   0.02362  -0.0007   1.0000   0.1530
  -4.750  -0.4729   0.03254   0.02150   0.0008   1.0000   0.1534
  -4.500  -0.4493   0.03049   0.01934   0.0018   1.0000   0.1559
  -4.250  -0.4248   0.02879   0.01767   0.0027   1.0000   0.1616
  -4.000  -0.3990   0.02747   0.01623   0.0037   1.0000   0.1664
  -3.750  -0.3716   0.02610   0.01487   0.0044   1.0000   0.1737
  -3.500  -0.3450   0.02502   0.01381   0.0053   1.0000   0.1843
  -3.250  -0.3202   0.02389   0.01283   0.0062   1.0000   0.1966
  -3.000  -0.3002   0.02291   0.01199   0.0076   1.0000   0.2150
  -2.500   0.0122   0.02174   0.01295  -0.0225   1.0000   1.0000
  -2.250   0.0208   0.02135   0.01252  -0.0209   1.0000   1.0000
  -2.000   0.0250   0.02108   0.01223  -0.0186   1.0000   1.0000
  -1.750   0.0232   0.02096   0.01211  -0.0154   1.0000   1.0000
  -1.500   0.0140   0.02101   0.01217  -0.0111   1.0000   1.0000
  -1.250  -0.0022   0.02123   0.01240  -0.0060   1.0000   1.0000
  -1.000  -0.0206   0.02159   0.01274  -0.0008   1.0000   1.0000
  -0.750  -0.0366   0.02205   0.01317   0.0037   1.0000   1.0000
  -0.500   0.0374   0.02264   0.01360  -0.0070   0.9773   1.0000
  -0.250   0.1472   0.02272   0.01356  -0.0227   0.9434   1.0000
   0.000   0.2516   0.02214   0.01293  -0.0361   0.9076   1.0000
   0.250   0.3069   0.02168   0.01242  -0.0398   0.8637   1.0000
   0.500   0.3414   0.02137   0.01198  -0.0393   0.8220   1.0000
   0.750   0.3626   0.02136   0.01182  -0.0366   0.7804   1.0000
   1.000   0.3821   0.02141   0.01168  -0.0336   0.7426   1.0000
   1.250   0.4011   0.02157   0.01164  -0.0307   0.7070   1.0000
   1.500   0.4206   0.02185   0.01173  -0.0283   0.6734   1.0000
   1.750   0.4407   0.02224   0.01194  -0.0263   0.6417   1.0000
   2.000   0.4611   0.02272   0.01229  -0.0246   0.6131   1.0000
   2.250   0.4826   0.02316   0.01253  -0.0228   0.5898   1.0000
   2.500   0.5032   0.02380   0.01309  -0.0215   0.5663   1.0000
   2.750   0.5244   0.02442   0.01364  -0.0202   0.5463   1.0000
   3.000   0.5458   0.02510   0.01423  -0.0190   0.5285   1.0000
   3.250   0.5669   0.02583   0.01491  -0.0178   0.5122   1.0000
   3.500   0.5879   0.02664   0.01568  -0.0167   0.4974   1.0000
   3.750   0.6090   0.02749   0.01652  -0.0156   0.4842   1.0000
   4.000   0.6311   0.02824   0.01717  -0.0144   0.4723   1.0000
   4.250   0.6506   0.02930   0.01832  -0.0133   0.4596   1.0000
   4.500   0.6695   0.03051   0.01963  -0.0123   0.4485   1.0000
   4.750   0.6916   0.03138   0.02044  -0.0112   0.4388   1.0000
   5.000   0.7084   0.03279   0.02202  -0.0101   0.4280   1.0000
   5.250   0.7266   0.03418   0.02349  -0.0090   0.4191   1.0000
   5.500   0.7452   0.03548   0.02488  -0.0079   0.4100   1.0000
   5.750   0.7594   0.03731   0.02688  -0.0068   0.4013   1.0000
   6.000   0.7784   0.03864   0.02825  -0.0056   0.3930   1.0000
   6.250   0.7885   0.04105   0.03087  -0.0045   0.3855   1.0000
   6.500   0.8035   0.04284   0.03277  -0.0034   0.3774   1.0000
   6.750   0.8123   0.04544   0.03555  -0.0023   0.3704   1.0000
   7.000   0.8123   0.04900   0.03935  -0.0015   0.3639   1.0000
   7.250   0.8372   0.05000   0.04032  -0.0004   0.3568   1.0000
   7.500   0.8037   0.05727   0.04791  -0.0004   0.3527   1.0000
   7.750   0.7568   0.06644   0.05719  -0.0026   0.3521   1.0000
   8.000   0.7171   0.07524   0.06594  -0.0059   0.3551   1.0000
   8.250   0.6978   0.08213   0.07282  -0.0088   0.3591   1.0000
   8.500   0.7025   0.08646   0.07718  -0.0095   0.3578   1.0000
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