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NACA 0010 (naca0010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010 (naca0010-il)
Reynolds number: 50,000
Max Cl/Cd: 25.9 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca0010-il-50000.txt
Download as CSV file: xf-naca0010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.7178   0.08934   0.08227  -0.0040   1.0000   0.1846
  -9.000  -0.7157   0.08339   0.07632  -0.0055   1.0000   0.1784
  -8.750  -0.7856   0.06961   0.06229  -0.0150   1.0000   0.1541
  -8.500  -0.7896   0.06406   0.05656  -0.0153   1.0000   0.1519
  -8.250  -0.7945   0.05868   0.05086  -0.0153   1.0000   0.1510
  -8.000  -0.7949   0.05380   0.04556  -0.0147   1.0000   0.1515
  -7.750  -0.7903   0.04929   0.04056  -0.0138   1.0000   0.1522
  -7.500  -0.7809   0.04514   0.03588  -0.0126   1.0000   0.1528
  -7.250  -0.7677   0.04149   0.03163  -0.0113   1.0000   0.1545
  -7.000  -0.7491   0.03874   0.02871  -0.0103   1.0000   0.1605
  -6.750  -0.7299   0.03631   0.02592  -0.0091   1.0000   0.1677
  -6.500  -0.7086   0.03364   0.02287  -0.0081   1.0000   0.1735
  -6.250  -0.6851   0.03148   0.02050  -0.0071   1.0000   0.1825
  -6.000  -0.6617   0.02956   0.01850  -0.0062   1.0000   0.1962
  -5.750  -0.6388   0.02779   0.01666  -0.0051   1.0000   0.2165
  -5.500  -0.6151   0.02600   0.01495  -0.0039   1.0000   0.2445
  -5.250  -0.5944   0.02439   0.01364  -0.0023   1.0000   0.2904
  -5.000  -0.5765   0.02294   0.01275  -0.0001   1.0000   0.3557
  -4.750  -0.5611   0.02183   0.01218   0.0030   1.0000   0.4370
  -4.500  -0.5458   0.02107   0.01186   0.0065   1.0000   0.5194
  -4.250  -0.5289   0.02061   0.01174   0.0104   1.0000   0.5921
  -4.000  -0.5113   0.02037   0.01171   0.0145   1.0000   0.6574
  -3.750  -0.4921   0.02038   0.01186   0.0188   1.0000   0.7159
  -3.500  -0.4692   0.02063   0.01213   0.0229   1.0000   0.7691
  -3.250  -0.4369   0.02110   0.01251   0.0255   1.0000   0.8201
  -3.000  -0.3860   0.02170   0.01286   0.0243   1.0000   0.8695
  -2.750  -0.2973   0.02226   0.01295   0.0151   1.0000   0.9154
  -2.500  -0.1955   0.02199   0.01223   0.0013   1.0000   0.9559
  -2.250  -0.1029   0.02098   0.01090  -0.0124   1.0000   0.9921
  -2.000  -0.0702   0.02006   0.00988  -0.0159   1.0000   1.0000
  -1.750  -0.0580   0.01940   0.00918  -0.0153   1.0000   1.0000
  -1.500  -0.0463   0.01884   0.00859  -0.0143   1.0000   1.0000
  -1.250  -0.0354   0.01839   0.00812  -0.0129   1.0000   1.0000
  -1.000  -0.0257   0.01803   0.00776  -0.0111   1.0000   1.0000
  -0.750  -0.0175   0.01776   0.00750  -0.0088   1.0000   1.0000
  -0.500  -0.0110   0.01757   0.00732  -0.0060   1.0000   1.0000
  -0.250  -0.0053   0.01745   0.00721  -0.0031   1.0000   1.0000
   0.000   0.0000   0.01742   0.00718   0.0000   1.0000   1.0000
   0.250   0.0053   0.01745   0.00721   0.0031   1.0000   1.0000
   0.500   0.0110   0.01757   0.00731   0.0060   1.0000   1.0000
   0.750   0.0176   0.01776   0.00750   0.0088   1.0000   1.0000
   1.000   0.0257   0.01803   0.00776   0.0111   1.0000   1.0000
   1.250   0.0354   0.01839   0.00812   0.0129   1.0000   1.0000
   1.500   0.0464   0.01884   0.00859   0.0143   1.0000   1.0000
   1.750   0.0581   0.01939   0.00917   0.0153   1.0000   1.0000
   2.000   0.0703   0.02006   0.00987   0.0159   1.0000   1.0000
   2.250   0.1029   0.02098   0.01089   0.0124   0.9921   1.0000
   2.500   0.1955   0.02199   0.01223  -0.0013   0.9559   1.0000
   2.750   0.2974   0.02226   0.01294  -0.0151   0.9155   1.0000
   3.000   0.3860   0.02170   0.01285  -0.0243   0.8695   1.0000
   3.250   0.4369   0.02110   0.01251  -0.0255   0.8201   1.0000
   3.500   0.4692   0.02063   0.01213  -0.0229   0.7692   1.0000
   3.750   0.4921   0.02038   0.01186  -0.0188   0.7159   1.0000
   4.000   0.5113   0.02037   0.01171  -0.0145   0.6574   1.0000
   4.250   0.5289   0.02061   0.01174  -0.0104   0.5921   1.0000
   4.500   0.5458   0.02107   0.01186  -0.0065   0.5194   1.0000
   4.750   0.5610   0.02183   0.01218  -0.0030   0.4370   1.0000
   5.000   0.5765   0.02294   0.01275   0.0001   0.3557   1.0000
   5.250   0.5944   0.02439   0.01364   0.0023   0.2904   1.0000
   5.500   0.6151   0.02599   0.01495   0.0039   0.2445   1.0000
   5.750   0.6388   0.02779   0.01666   0.0051   0.2165   1.0000
   6.000   0.6617   0.02956   0.01850   0.0062   0.1962   1.0000
   6.250   0.6851   0.03148   0.02050   0.0071   0.1825   1.0000
   6.500   0.7086   0.03364   0.02287   0.0081   0.1736   1.0000
   6.750   0.7299   0.03630   0.02591   0.0091   0.1677   1.0000
   7.000   0.7492   0.03874   0.02871   0.0103   0.1605   1.0000
   7.250   0.7678   0.04149   0.03164   0.0113   0.1545   1.0000
   7.500   0.7809   0.04514   0.03588   0.0126   0.1528   1.0000
   7.750   0.7904   0.04929   0.04057   0.0138   0.1522   1.0000
   8.000   0.7949   0.05380   0.04556   0.0147   0.1515   1.0000
   8.250   0.7946   0.05869   0.05086   0.0153   0.1510   1.0000
   8.500   0.7897   0.06408   0.05657   0.0153   0.1519   1.0000
   8.750   0.7857   0.06963   0.06231   0.0149   0.1541   1.0000
   9.000   0.7159   0.08346   0.07640   0.0054   0.1786   1.0000
   9.250   0.7090   0.09293   0.08587  -0.0001   0.2011   1.0000
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