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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 500,000
Max Cl/Cd: 86.39 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa46-il-500000.txt
Download as CSV file: xf-usa46-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 46 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5418   0.06789   0.06579  -0.0256   1.0000   0.0168
  -7.750  -0.5931   0.02848   0.02525  -0.0570   1.0000   0.0122
  -7.500  -0.5766   0.02440   0.02067  -0.0578   1.0000   0.0124
  -7.250  -0.5537   0.02386   0.02005  -0.0575   1.0000   0.0134
  -7.000  -0.5330   0.02146   0.01729  -0.0573   1.0000   0.0142
  -6.750  -0.5119   0.01880   0.01421  -0.0569   1.0000   0.0147
  -6.500  -0.4893   0.01699   0.01208  -0.0563   1.0000   0.0154
  -6.250  -0.4657   0.01585   0.01070  -0.0556   1.0000   0.0160
  -6.000  -0.4438   0.01353   0.00804  -0.0549   1.0000   0.0173
  -5.750  -0.4201   0.01274   0.00717  -0.0542   1.0000   0.0190
  -5.500  -0.3959   0.01241   0.00678  -0.0535   1.0000   0.0213
  -5.250  -0.3716   0.01204   0.00632  -0.0527   1.0000   0.0232
  -5.000  -0.3483   0.01091   0.00510  -0.0519   1.0000   0.0274
  -4.750  -0.3242   0.01064   0.00478  -0.0511   1.0000   0.0318
  -4.500  -0.3007   0.01014   0.00427  -0.0504   1.0000   0.0389
  -4.250  -0.2772   0.01007   0.00413  -0.0495   1.0000   0.0436
  -4.000  -0.2541   0.00954   0.00360  -0.0487   1.0000   0.0502
  -3.750  -0.2238   0.00942   0.00343  -0.0493   0.9986   0.0557
  -3.500  -0.1870   0.00901   0.00304  -0.0513   0.9955   0.0647
  -3.250  -0.1509   0.00870   0.00271  -0.0531   0.9920   0.0751
  -3.000  -0.1152   0.00836   0.00244  -0.0548   0.9880   0.0917
  -2.750  -0.0771   0.00811   0.00231  -0.0570   0.9842   0.1237
  -2.500  -0.0403   0.00795   0.00219  -0.0589   0.9782   0.1471
  -2.250  -0.0017   0.00781   0.00205  -0.0611   0.9724   0.1636
  -2.000   0.0349   0.00762   0.00190  -0.0628   0.9659   0.1766
  -1.750   0.0712   0.00742   0.00173  -0.0645   0.9579   0.1888
  -1.500   0.1060   0.00723   0.00158  -0.0659   0.9470   0.2036
  -1.250   0.1408   0.00702   0.00143  -0.0672   0.9342   0.2232
  -1.000   0.1736   0.00674   0.00131  -0.0682   0.9182   0.2669
  -0.750   0.2018   0.00610   0.00124  -0.0684   0.8995   0.4612
  -0.500   0.2251   0.00553   0.00132  -0.0673   0.8795   0.6744
  -0.250   0.2496   0.00542   0.00136  -0.0661   0.8603   0.7621
   0.000   0.2747   0.00538   0.00135  -0.0651   0.8407   0.8011
   0.250   0.2993   0.00529   0.00131  -0.0641   0.8203   0.8362
   0.500   0.3274   0.00507   0.00126  -0.0636   0.7990   0.9109
   0.750   0.3667   0.00512   0.00123  -0.0659   0.7737   1.0000
   1.000   0.3927   0.00525   0.00123  -0.0653   0.7460   1.0000
   1.250   0.4185   0.00541   0.00126  -0.0647   0.7161   1.0000
   1.500   0.4445   0.00557   0.00129  -0.0641   0.6862   1.0000
   1.750   0.4706   0.00574   0.00135  -0.0636   0.6576   1.0000
   2.000   0.4965   0.00593   0.00141  -0.0630   0.6267   1.0000
   2.250   0.5219   0.00617   0.00151  -0.0624   0.5907   1.0000
   2.500   0.5477   0.00640   0.00161  -0.0619   0.5575   1.0000
   2.750   0.5736   0.00664   0.00173  -0.0614   0.5253   1.0000
   3.000   0.5986   0.00695   0.00186  -0.0608   0.4763   1.0000
   3.250   0.6231   0.00736   0.00201  -0.0602   0.4051   1.0000
   3.500   0.6477   0.00781   0.00221  -0.0596   0.3491   1.0000
   3.750   0.6726   0.00826   0.00245  -0.0591   0.3037   1.0000
   4.000   0.6977   0.00868   0.00270  -0.0586   0.2594   1.0000
   4.250   0.7206   0.00947   0.00305  -0.0579   0.1583   1.0000
   4.500   0.7397   0.01104   0.00393  -0.0567   0.0293   1.0000
   4.750   0.7648   0.01159   0.00449  -0.0561   0.0223   1.0000
   5.000   0.7895   0.01223   0.00523  -0.0553   0.0201   1.0000
   5.250   0.8141   0.01284   0.00593  -0.0546   0.0186   1.0000
   5.500   0.8379   0.01358   0.00676  -0.0537   0.0174   1.0000
   5.750   0.8614   0.01434   0.00756  -0.0529   0.0159   1.0000
   6.000   0.8820   0.01567   0.00897  -0.0517   0.0146   1.0000
   6.250   0.9001   0.01800   0.01143  -0.0500   0.0139   1.0000
   6.500   0.9245   0.01878   0.01231  -0.0492   0.0135   1.0000
   6.750   0.9476   0.02012   0.01379  -0.0482   0.0132   1.0000
   7.000   0.9700   0.02205   0.01591  -0.0471   0.0131   1.0000
   7.250   0.9907   0.02504   0.01917  -0.0457   0.0134   1.0000
   7.500   1.0150   0.02663   0.02087  -0.0447   0.0148   1.0000
   7.750   1.0334   0.03363   0.02866  -0.0413   0.0229   1.0000
   8.000   1.0470   0.03760   0.03278  -0.0404   0.0216   1.0000
   8.250   1.0520   0.04088   0.03668  -0.0377   0.0194   1.0000
   8.500   1.0650   0.04321   0.03931  -0.0360   0.0179   1.0000
   8.750   1.0731   0.04611   0.04248  -0.0343   0.0168   1.0000
   9.000   1.0783   0.04911   0.04569  -0.0328   0.0161   1.0000
   9.250   1.0814   0.05205   0.04879  -0.0313   0.0156   1.0000
   9.500   1.0826   0.05500   0.05187  -0.0300   0.0152   1.0000
   9.750   1.0803   0.05819   0.05517  -0.0286   0.0149   1.0000
  10.000   1.0667   0.06151   0.05865  -0.0263   0.0148   1.0000
  10.250   1.0476   0.06486   0.06213  -0.0244   0.0147   1.0000
  10.500   1.0274   0.06895   0.06636  -0.0247   0.0147   1.0000
  10.750   1.0065   0.07418   0.07174  -0.0273   0.0147   1.0000
  11.000   0.9607   0.08718   0.08503  -0.0388   0.0154   1.0000
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