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NACA 64(1)-212 (n64212-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 64(1)-212 (n64212-il)
Reynolds number: 50,000
Max Cl/Cd: 31.06 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n64212-il-50000.txt
Download as CSV file: xf-n64212-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64(1)-212                                  
  
 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  -0.4470   0.10120   0.09406  -0.0029   1.0000   0.3884
  -8.750  -0.5415   0.08244   0.07565  -0.0288   1.0000   0.2004
  -8.500  -0.6112   0.07267   0.06603  -0.0383   1.0000   0.1791
  -8.250  -0.6327   0.06585   0.05900  -0.0410   1.0000   0.1591
  -8.000  -0.6557   0.06033   0.05300  -0.0415   1.0000   0.1463
  -7.750  -0.6484   0.05577   0.04835  -0.0408   1.0000   0.1400
  -7.500  -0.6572   0.05171   0.04354  -0.0392   1.0000   0.1318
  -7.250  -0.6477   0.04810   0.03976  -0.0377   1.0000   0.1293
  -7.000  -0.6402   0.04501   0.03633  -0.0359   1.0000   0.1284
  -6.750  -0.6313   0.04227   0.03320  -0.0339   1.0000   0.1288
  -6.500  -0.6200   0.03974   0.03026  -0.0319   1.0000   0.1296
  -6.250  -0.6057   0.03732   0.02747  -0.0301   1.0000   0.1300
  -6.000  -0.5895   0.03510   0.02487  -0.0283   1.0000   0.1309
  -5.750  -0.5722   0.03313   0.02254  -0.0265   1.0000   0.1337
  -5.500  -0.5539   0.03128   0.02076  -0.0251   1.0000   0.1406
  -5.250  -0.5341   0.02981   0.01899  -0.0234   1.0000   0.1467
  -5.000  -0.5134   0.02820   0.01749  -0.0218   1.0000   0.1547
  -4.750  -0.4945   0.02693   0.01628  -0.0200   1.0000   0.1686
  -4.500  -0.4769   0.02571   0.01516  -0.0179   1.0000   0.1851
  -4.250  -0.4621   0.02433   0.01401  -0.0159   1.0000   0.2141
  -4.000  -0.4517   0.02162   0.01248  -0.0142   1.0000   0.3159
  -3.750  -0.4663   0.02211   0.01481  -0.0015   1.0000   0.6655
  -3.500  -0.4693   0.02360   0.01627   0.0083   1.0000   0.7301
  -3.250  -0.4684   0.02465   0.01723   0.0167   1.0000   0.7735
  -3.000  -0.3029   0.02990   0.02159   0.0116   1.0000   0.9143
  -2.750  -0.2317   0.02928   0.02055   0.0028   1.0000   0.9366
  -2.500  -0.1922   0.02869   0.01972  -0.0014   1.0000   0.9521
  -2.250  -0.1548   0.02816   0.01900  -0.0055   1.0000   0.9657
  -2.000  -0.1163   0.02769   0.01834  -0.0101   1.0000   0.9786
  -1.750  -0.0758   0.02727   0.01778  -0.0152   1.0000   0.9911
  -1.500  -0.0463   0.02703   0.01743  -0.0185   1.0000   1.0000
  -1.250  -0.0535   0.02701   0.01739  -0.0154   1.0000   1.0000
  -1.000  -0.0608   0.02693   0.01729  -0.0124   1.0000   1.0000
  -0.750  -0.0682   0.02679   0.01713  -0.0093   1.0000   1.0000
  -0.500  -0.0761   0.02658   0.01689  -0.0062   1.0000   1.0000
  -0.250  -0.0845   0.02630   0.01658  -0.0030   1.0000   1.0000
   0.000  -0.0937   0.02593   0.01618   0.0004   1.0000   1.0000
   0.250  -0.1030   0.02549   0.01570   0.0038   1.0000   1.0000
   0.500  -0.0834   0.02566   0.01580   0.0023   0.9934   1.0000
   0.750  -0.0612   0.02596   0.01603   0.0004   0.9853   1.0000
   1.000  -0.0387   0.02635   0.01632  -0.0012   0.9780   1.0000
   1.250  -0.0071   0.02709   0.01697  -0.0043   0.9708   1.0000
   1.500   0.0175   0.02762   0.01744  -0.0060   0.9633   1.0000
   1.750   0.0543   0.02864   0.01839  -0.0096   0.9562   1.0000
   2.000   0.0772   0.02923   0.01893  -0.0108   0.9485   1.0000
   2.250   0.1135   0.03031   0.01999  -0.0142   0.9405   1.0000
   2.500   0.1377   0.03105   0.02073  -0.0154   0.9322   1.0000
   2.750   0.1701   0.03211   0.02179  -0.0179   0.9236   1.0000
   3.000   0.1988   0.03308   0.02279  -0.0198   0.9147   1.0000
   3.250   0.2247   0.03406   0.02381  -0.0212   0.9056   1.0000
   3.500   0.2606   0.03530   0.02514  -0.0240   0.8953   1.0000
   3.750   0.2854   0.03627   0.02618  -0.0251   0.8844   1.0000
   4.000   0.3110   0.03734   0.02734  -0.0263   0.8727   1.0000
   4.250   0.3409   0.03850   0.02863  -0.0279   0.8592   1.0000
   4.500   0.3721   0.03963   0.02991  -0.0296   0.8438   1.0000
   4.750   0.4046   0.04073   0.03118  -0.0313   0.8267   1.0000
   5.000   0.4471   0.04175   0.03246  -0.0338   0.8067   1.0000
   5.250   0.4778   0.04257   0.03349  -0.0345   0.7842   1.0000
   5.500   0.5274   0.04280   0.03406  -0.0364   0.7561   1.0000
   5.750   0.6074   0.04016   0.03202  -0.0375   0.7085   1.0000
   6.000   0.7482   0.02598   0.01881  -0.0290   0.6005   1.0000
   6.250   0.7421   0.02389   0.01503  -0.0161   0.3133   1.0000
   6.500   0.7366   0.02714   0.01697  -0.0118   0.2226   1.0000
   6.750   0.7538   0.02944   0.01882  -0.0101   0.1818   1.0000
   7.000   0.7897   0.03162   0.02084  -0.0102   0.1565   1.0000
   7.250   0.8265   0.03388   0.02299  -0.0107   0.1394   1.0000
   7.500   0.8627   0.03661   0.02583  -0.0112   0.1307   1.0000
   7.750   0.8918   0.03919   0.02849  -0.0113   0.1226   1.0000
   8.000   0.9159   0.04226   0.03190  -0.0107   0.1188   1.0000
   8.250   0.9363   0.04547   0.03559  -0.0096   0.1176   1.0000
   8.500   0.9520   0.04885   0.03945  -0.0083   0.1167   1.0000
   8.750   0.9629   0.05233   0.04340  -0.0068   0.1158   1.0000
   9.000   0.9694   0.05600   0.04751  -0.0053   0.1152   1.0000
   9.250   0.9711   0.05999   0.05192  -0.0037   0.1159   1.0000
   9.500   0.9706   0.06438   0.05664  -0.0023   0.1177   1.0000
   9.750   0.9764   0.06930   0.06176  -0.0017   0.1199   1.0000
  10.000   0.9461   0.07338   0.06635   0.0004   0.1240   1.0000
  10.250   0.9040   0.07814   0.07135   0.0016   0.1271   1.0000
  10.500   0.8736   0.08388   0.07720   0.0004   0.1304   1.0000
  10.750   0.8768   0.08979   0.08318  -0.0007   0.1361   1.0000
  11.000   0.7905   0.10466   0.09804  -0.0137   0.1518   1.0000
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