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GOE 419 AIRFOIL (goe419-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 419 AIRFOIL (goe419-il)
Reynolds number: 500,000
Max Cl/Cd: 109.2 at α=1.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe419-il-500000.txt
Download as CSV file: xf-goe419-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 419 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3789   0.08971   0.08759  -0.0205   1.0000   0.0100
  -7.500  -0.3823   0.08759   0.08552  -0.0196   1.0000   0.0102
  -7.250  -0.3847   0.08524   0.08321  -0.0192   1.0000   0.0103
  -7.000  -0.3831   0.08253   0.08053  -0.0199   1.0000   0.0105
  -6.750  -0.3826   0.07994   0.07798  -0.0203   1.0000   0.0108
  -6.500  -0.3787   0.07723   0.07530  -0.0216   0.9994   0.0111
  -6.250  -0.3477   0.07226   0.07029  -0.0299   0.9962   0.0116
  -6.000  -0.3154   0.06733   0.06532  -0.0380   0.9928   0.0125
  -5.750  -0.2806   0.06240   0.06033  -0.0461   0.9883   0.0137
  -5.500  -0.2333   0.05803   0.05581  -0.0557   0.9847   0.0151
  -5.250  -0.1940   0.05362   0.05130  -0.0624   0.9819   0.0154
  -5.000  -0.1618   0.04932   0.04689  -0.0670   0.9754   0.0155
  -4.750  -0.1251   0.04464   0.04207  -0.0722   0.9718   0.0156
  -4.500  -0.0915   0.04010   0.03737  -0.0760   0.9663   0.0156
  -4.250  -0.0595   0.03224   0.02922  -0.0806   0.9600   0.0162
  -4.000  -0.0288   0.02790   0.02464  -0.0833   0.9545   0.0173
  -3.000   0.1108   0.01269   0.00752  -0.0890   0.9316   0.0165
  -2.750   0.1453   0.01040   0.00484  -0.0901   0.9237   0.0178
  -2.500   0.1829   0.00969   0.00396  -0.0921   0.9160   0.0222
  -2.250   0.2175   0.00937   0.00352  -0.0935   0.9071   0.0273
  -2.000   0.2474   0.00860   0.00265  -0.0939   0.8970   0.0367
  -1.750   0.2761   0.00822   0.00218  -0.0941   0.8867   0.0449
  -1.500   0.3046   0.00808   0.00197  -0.0942   0.8767   0.0501
  -1.250   0.3314   0.00793   0.00172  -0.0939   0.8661   0.0524
  -1.000   0.3571   0.00771   0.00148  -0.0934   0.8553   0.0621
  -0.750   0.3821   0.00737   0.00133  -0.0928   0.8447   0.1342
  -0.500   0.4068   0.00708   0.00133  -0.0923   0.8336   0.2560
  -0.250   0.4742   0.00532   0.00136  -0.1017   0.8257   1.0000
   0.000   0.4982   0.00538   0.00132  -0.1008   0.8129   1.0000
   0.250   0.5221   0.00545   0.00130  -0.0999   0.7995   1.0000
   0.500   0.5460   0.00551   0.00129  -0.0989   0.7850   1.0000
   0.750   0.5696   0.00558   0.00129  -0.0980   0.7690   1.0000
   1.000   0.5932   0.00566   0.00130  -0.0970   0.7511   1.0000
   1.250   0.6167   0.00575   0.00134  -0.0960   0.7319   1.0000
   1.500   0.6399   0.00586   0.00138  -0.0949   0.7100   1.0000
   1.750   0.6588   0.00613   0.00139  -0.0929   0.6607   1.0000
   2.000   0.6796   0.00641   0.00149  -0.0913   0.6257   1.0000
   2.250   0.6997   0.00673   0.00161  -0.0897   0.5805   1.0000
   2.500   0.7145   0.00734   0.00173  -0.0870   0.4694   1.0000
   2.750   0.7136   0.00966   0.00245  -0.0820   0.1491   1.0000
   3.000   0.7296   0.01080   0.00314  -0.0798   0.0234   1.0000
   3.250   0.7528   0.01122   0.00376  -0.0787   0.0199   1.0000
   3.500   0.7753   0.01170   0.00435  -0.0776   0.0186   1.0000
   3.750   0.7974   0.01222   0.00492  -0.0764   0.0159   1.0000
   4.000   0.8176   0.01296   0.00574  -0.0748   0.0141   1.0000
   4.250   0.8358   0.01391   0.00677  -0.0728   0.0131   1.0000
   4.500   0.8533   0.01502   0.00796  -0.0706   0.0134   1.0000
   4.750   0.8715   0.01649   0.00952  -0.0682   0.0160   1.0000
   5.750   0.9680   0.02274   0.01614  -0.0640   0.0157   1.0000
   6.000   0.9906   0.02442   0.01792  -0.0631   0.0135   1.0000
   6.250   1.0125   0.02740   0.02094  -0.0625   0.0116   1.0000
   6.500   1.0249   0.03556   0.02959  -0.0601   0.0103   1.0000
   6.750   1.0396   0.03842   0.03274  -0.0578   0.0103   1.0000
   7.000   1.0522   0.04133   0.03595  -0.0553   0.0103   1.0000
   7.250   1.0631   0.04414   0.03905  -0.0528   0.0103   1.0000
   7.500   1.0717   0.04698   0.04218  -0.0501   0.0102   1.0000
   7.750   1.0787   0.04970   0.04523  -0.0473   0.0101   1.0000
   8.000   1.0847   0.05198   0.04778  -0.0445   0.0100   1.0000
   8.250   1.0990   0.05196   0.04798  -0.0420   0.0093   1.0000
   8.500   1.1073   0.05373   0.04999  -0.0391   0.0085   1.0000
   8.750   1.1076   0.05660   0.05309  -0.0362   0.0081   1.0000
   9.000   1.1030   0.05969   0.05639  -0.0331   0.0078   1.0000
   9.250   1.0932   0.06261   0.05948  -0.0297   0.0077   1.0000
   9.500   1.0778   0.06530   0.06232  -0.0258   0.0076   1.0000
   9.750   1.0602   0.06842   0.06557  -0.0231   0.0075   1.0000
  10.000   1.0421   0.07198   0.06926  -0.0219   0.0075   1.0000
  10.250   1.0227   0.07624   0.07365  -0.0221   0.0076   1.0000
  10.500   1.0031   0.08125   0.07878  -0.0238   0.0076   1.0000
  10.750   0.9837   0.08714   0.08477  -0.0269   0.0078   1.0000
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