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NACA 6412 (naca6412-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 6412 (naca6412-il)
Reynolds number: 50,000
Max Cl/Cd: 9.81 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca6412-il-50000.txt
Download as CSV file: xf-naca6412-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 6412                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3304   0.13030   0.12383  -0.0292   1.0001   0.1748
  -8.750  -0.3583   0.13182   0.12550  -0.0274   1.0001   0.1754
  -8.500  -0.3148   0.12222   0.11582  -0.0254   1.0001   0.1826
  -8.250  -0.3221   0.12090   0.11458  -0.0236   1.0001   0.1873
  -8.000  -0.3420   0.12106   0.11486  -0.0218   1.0001   0.1909
  -7.750  -0.3716   0.12225   0.11619  -0.0194   1.0001   0.1923
  -7.500  -0.4036   0.12347   0.11758  -0.0172   1.0001   0.1930
  -7.250  -0.3607   0.11491   0.10895  -0.0157   1.0001   0.1997
  -7.000  -0.3705   0.11359   0.10771  -0.0133   1.0001   0.2039
  -6.750  -0.3921   0.11331   0.10754  -0.0107   1.0001   0.2075
  -6.500  -0.4215   0.11416   0.10853  -0.0116   1.0001   0.2104
  -6.250  -0.4238   0.11077   0.10522  -0.0102   1.0001   0.2128
  -6.000  -0.4156   0.10733   0.10179  -0.0065   1.0001   0.2177
  -5.750  -0.4218   0.10560   0.10013  -0.0058   1.0001   0.2238
  -5.500  -0.4384   0.10518   0.09978  -0.0126   1.0001   0.2297
  -5.250  -0.4318   0.10107   0.09572  -0.0059   1.0001   0.2341
  -5.000  -0.4327   0.09919   0.09386  -0.0060   1.0001   0.2435
  -4.750  -0.4335   0.09652   0.09123  -0.0079   1.0001   0.2501
  -4.500  -0.4305   0.09419   0.08894  -0.0058   1.0001   0.2592
  -4.250  -0.4270   0.09151   0.08628  -0.0084   1.0001   0.2687
  -4.000  -0.4173   0.08945   0.08418  -0.0133   1.0001   0.2837
  -3.750  -0.3873   0.08592   0.08059  -0.0182   0.9930   0.3015
  -3.500  -0.3563   0.08220   0.07685  -0.0206   0.9829   0.3215
  -3.250  -0.3242   0.07938   0.07395  -0.0259   0.9723   0.3544
  -3.000  -0.3042   0.07643   0.07101  -0.0252   0.9622   0.3773
  -2.750  -0.2800   0.07380   0.06835  -0.0265   0.9523   0.4126
  -2.500  -0.2622   0.07155   0.06613  -0.0247   0.9436   0.4643
  -2.250  -0.0529   0.05589   0.04808  -0.0876   0.9317   0.2070
  -2.000  -0.0005   0.05297   0.04454  -0.0944   0.9228   0.2079
  -1.750   0.0439   0.05090   0.04195  -0.0990   0.9135   0.2138
  -1.500   0.0783   0.04999   0.04076  -0.1014   0.9041   0.2246
  -1.250   0.1235   0.04895   0.03939  -0.1053   0.8955   0.2381
  -1.000   0.1507   0.04848   0.03863  -0.1063   0.8858   0.2517
  -0.750   0.1967   0.04802   0.03794  -0.1099   0.8781   0.2739
  -0.500   0.2151   0.04806   0.03785  -0.1095   0.8685   0.2907
  -0.250   0.2630   0.04779   0.03739  -0.1132   0.8611   0.3230
   0.000   0.2776   0.04812   0.03764  -0.1121   0.8515   0.3451
   0.250   0.3266   0.04795   0.03740  -0.1158   0.8445   0.3941
   0.500   0.3387   0.04843   0.03797  -0.1145   0.8354   0.4291
   0.750   0.3841   0.04803   0.03809  -0.1175   0.8286   0.5310
   1.000   0.3871   0.04738   0.03867  -0.1139   0.8205   0.9999
   1.250   0.4254   0.04852   0.03904  -0.1163   0.8130   0.9999
   1.500   0.4386   0.04991   0.04017  -0.1154   0.8050   0.9999
   1.750   0.4645   0.05112   0.04112  -0.1160   0.7976   0.9999
   2.000   0.4843   0.05252   0.04233  -0.1159   0.7905   0.9999
   2.250   0.4988   0.05398   0.04366  -0.1153   0.7833   0.9999
   2.500   0.5342   0.05516   0.04465  -0.1170   0.7765   0.9999
   2.750   0.5333   0.05703   0.04647  -0.1147   0.7697   0.9999
   3.000   0.5632   0.05831   0.04763  -0.1158   0.7628   0.9999
   3.250   0.5723   0.06014   0.04939  -0.1148   0.7564   0.9999
   3.500   0.5833   0.06193   0.05113  -0.1140   0.7502   0.9999
   3.750   0.6202   0.06320   0.05230  -0.1157   0.7431   0.9999
   4.000   0.6139   0.06554   0.05465  -0.1134   0.7382   0.9999
   4.250   0.6273   0.06743   0.05651  -0.1130   0.7325   0.9999
   4.500   0.6555   0.06903   0.05806  -0.1139   0.7252   0.9999
   4.750   0.6527   0.07151   0.06055  -0.1123   0.7217   0.9999
   5.000   0.6602   0.07377   0.06282  -0.1116   0.7175   0.9999
   5.250   0.6929   0.07535   0.06437  -0.1128   0.7086   0.9999
   5.500   0.6905   0.07809   0.06714  -0.1115   0.7067   0.9999
   5.750   0.6933   0.08088   0.06996  -0.1109   0.7062   0.9999
   6.000   0.7021   0.08387   0.07298  -0.1112   0.7080   0.9999
   6.250   0.7197   0.08705   0.07620  -0.1123   0.7101   0.9999
   6.500   0.6387   0.09319   0.08249  -0.1086   0.7826   0.9999
   6.750   0.6586   0.09609   0.08541  -0.1095   0.7752   0.9999
   7.000   0.6765   0.09828   0.08764  -0.1099   0.7629   0.9999
   7.250   0.6788   0.10023   0.08962  -0.1086   0.7534   0.9999
   7.500   0.7178   0.10418   0.09361  -0.1117   0.7437   0.9999
   7.750   0.7125   0.10520   0.09468  -0.1093   0.7310   0.9999
   8.000   0.7210   0.10767   0.09719  -0.1089   0.7213   0.9999
   8.250   0.7611   0.11176   0.10136  -0.1119   0.7102   0.9999
   8.500   0.7524   0.11275   0.10240  -0.1094   0.6972   0.9999
   8.750   0.7588   0.11523   0.10494  -0.1088   0.6864   0.9999
   9.000   0.7932   0.11926   0.10905  -0.1111   0.6760   0.9999
   9.250   0.7950   0.12088   0.11076  -0.1098   0.6619   0.9999
   9.500   0.7958   0.12308   0.11302  -0.1088   0.6495   0.9999
   9.750   0.8108   0.12628   0.11631  -0.1093   0.6392   0.9999
  10.000   0.8434   0.13017   0.12030  -0.1109   0.6255   0.9999
  10.250   0.8399   0.13173   0.12194  -0.1095   0.6113   0.9999
  10.500   0.8408   0.13412   0.12442  -0.1088   0.5981   0.9999
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