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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 200,000
Max Cl/Cd: 48.45 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa49-il-200000-n5.txt
Download as CSV file: xf-usa49-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 49 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4414   0.09547   0.09220  -0.0097   1.0000   0.0225
  -9.000  -0.4422   0.09112   0.08787  -0.0111   1.0000   0.0225
  -6.000  -0.4801   0.04251   0.03830  -0.0321   1.0000   0.0115
  -5.750  -0.4671   0.03748   0.03292  -0.0321   1.0000   0.0111
  -5.500  -0.4518   0.03310   0.02814  -0.0315   1.0000   0.0108
  -5.250  -0.4346   0.02936   0.02395  -0.0304   1.0000   0.0106
  -5.000  -0.4158   0.02624   0.02037  -0.0292   1.0000   0.0106
  -4.750  -0.3953   0.02365   0.01731  -0.0279   1.0000   0.0107
  -4.500  -0.3735   0.02178   0.01506  -0.0267   1.0000   0.0111
  -4.250  -0.3505   0.02125   0.01429  -0.0255   1.0000   0.0122
  -4.000  -0.3285   0.01835   0.01095  -0.0244   1.0000   0.0129
  -3.750  -0.3054   0.01669   0.00908  -0.0233   1.0000   0.0133
  -3.500  -0.2821   0.01537   0.00764  -0.0224   1.0000   0.0140
  -3.250  -0.2563   0.01434   0.00651  -0.0220   0.9992   0.0150
  -3.000  -0.2218   0.01355   0.00564  -0.0235   0.9955   0.0180
  -2.750  -0.1872   0.01285   0.00480  -0.0250   0.9910   0.0204
  -2.500  -0.1512   0.01228   0.00409  -0.0266   0.9870   0.0235
  -2.250  -0.1170   0.01160   0.00345  -0.0281   0.9818   0.0425
  -2.000  -0.0820   0.01083   0.00318  -0.0301   0.9774   0.1615
  -1.750  -0.0505   0.01010   0.00295  -0.0314   0.9703   0.2920
  -1.500  -0.0252   0.00856   0.00287  -0.0313   0.9637   0.6650
  -1.250   0.0008   0.00811   0.00289  -0.0300   0.9556   0.8150
  -1.000   0.0327   0.00801   0.00284  -0.0302   0.9464   0.8965
  -0.750   0.0755   0.00795   0.00271  -0.0329   0.9370   0.9482
  -0.500   0.1226   0.00788   0.00256  -0.0367   0.9236   0.9883
  -0.250   0.1592   0.00783   0.00240  -0.0384   0.9055   1.0000
   0.000   0.1859   0.00780   0.00228  -0.0379   0.8854   1.0000
   0.250   0.2115   0.00778   0.00218  -0.0372   0.8626   1.0000
   0.500   0.2367   0.00779   0.00209  -0.0364   0.8371   1.0000
   0.750   0.2612   0.00782   0.00198  -0.0353   0.8028   1.0000
   1.000   0.2849   0.00793   0.00187  -0.0340   0.7565   1.0000
   1.250   0.3082   0.00814   0.00180  -0.0327   0.7007   1.0000
   1.500   0.3318   0.00842   0.00179  -0.0316   0.6440   1.0000
   1.750   0.3559   0.00873   0.00181  -0.0308   0.5908   1.0000
   2.000   0.3807   0.00901   0.00188  -0.0302   0.5497   1.0000
   2.250   0.4061   0.00927   0.00200  -0.0297   0.5158   1.0000
   2.500   0.4314   0.00955   0.00211  -0.0293   0.4736   1.0000
   2.750   0.4566   0.00987   0.00224  -0.0288   0.4254   1.0000
   3.000   0.4811   0.01031   0.00241  -0.0284   0.3595   1.0000
   3.250   0.5056   0.01082   0.00263  -0.0280   0.2964   1.0000
   3.500   0.5301   0.01138   0.00291  -0.0277   0.2305   1.0000
   3.750   0.5544   0.01203   0.00324  -0.0274   0.1614   1.0000
   4.000   0.5774   0.01298   0.00377  -0.0270   0.0766   1.0000
   4.250   0.6032   0.01337   0.00418  -0.0267   0.0706   1.0000
   4.500   0.6288   0.01379   0.00465  -0.0264   0.0665   1.0000
   4.750   0.6542   0.01427   0.00520  -0.0261   0.0635   1.0000
   5.000   0.6799   0.01467   0.00575  -0.0258   0.0621   1.0000
   5.250   0.7052   0.01514   0.00634  -0.0254   0.0607   1.0000
   5.500   0.7300   0.01568   0.00699  -0.0250   0.0595   1.0000
   5.750   0.7544   0.01629   0.00772  -0.0245   0.0584   1.0000
   6.000   0.7786   0.01691   0.00845  -0.0241   0.0548   1.0000
   6.250   0.8024   0.01752   0.00913  -0.0237   0.0477   1.0000
   6.500   0.8282   0.01778   0.00950  -0.0236   0.0415   1.0000
   6.750   0.8543   0.01800   0.00986  -0.0234   0.0340   1.0000
   7.000   0.8809   0.01818   0.01011  -0.0233   0.0227   1.0000
   7.250   0.9042   0.01896   0.01089  -0.0226   0.0168   1.0000
   7.500   0.9256   0.02007   0.01211  -0.0218   0.0131   1.0000
   7.750   0.9430   0.02205   0.01434  -0.0203   0.0113   1.0000
   8.000   0.9626   0.02367   0.01619  -0.0191   0.0107   1.0000
   8.250   0.9823   0.02531   0.01808  -0.0180   0.0098   1.0000
   8.500   1.0022   0.02672   0.01971  -0.0171   0.0088   1.0000
   8.750   1.0207   0.02828   0.02148  -0.0161   0.0080   1.0000
   9.000   1.0363   0.03042   0.02389  -0.0149   0.0076   1.0000
   9.250   1.0477   0.03327   0.02709  -0.0133   0.0073   1.0000
   9.750   1.0585   0.04041   0.03506  -0.0095   0.0070   1.0000
  10.000   1.0581   0.04411   0.03918  -0.0075   0.0069   1.0000
  10.250   1.0515   0.04801   0.04345  -0.0055   0.0069   1.0000
  10.500   1.0380   0.05168   0.04740  -0.0033   0.0069   1.0000
  10.750   1.0217   0.05590   0.05188  -0.0028   0.0069   1.0000
  11.000   1.0045   0.06095   0.05715  -0.0044   0.0069   1.0000
  11.250   0.9873   0.06718   0.06358  -0.0085   0.0069   1.0000
  11.500   0.9699   0.07482   0.07138  -0.0145   0.0070   1.0000
  11.750   0.9521   0.08378   0.08047  -0.0215   0.0071   1.0000
  12.000   0.9352   0.09364   0.09044  -0.0285   0.0071   1.0000
  12.250   0.8940   0.11329   0.11017  -0.0381   0.0078   1.0000
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