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E176 (8.83%) (e176-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: E176 (8.83%) (e176-il)
Reynolds number: 1,000,000
Max Cl/Cd: 132.94 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e176-il-1000000.txt
Download as CSV file: xf-e176-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E176  (8.83%)                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.3436   0.00484   0.00105  -0.0606   0.6051   0.9782
   0.500   0.3820   0.00489   0.00102  -0.0630   0.5998   1.0000
   0.750   0.4095   0.00492   0.00102  -0.0628   0.5944   1.0000
   1.000   0.4368   0.00498   0.00103  -0.0627   0.5887   1.0000
   1.250   0.4643   0.00503   0.00104  -0.0625   0.5825   1.0000
   1.500   0.4918   0.00508   0.00106  -0.0624   0.5756   1.0000
   1.750   0.5193   0.00514   0.00109  -0.0624   0.5691   1.0000
   2.000   0.5470   0.00520   0.00112  -0.0623   0.5615   1.0000
   2.250   0.5746   0.00527   0.00115  -0.0622   0.5541   1.0000
   2.500   0.6022   0.00534   0.00119  -0.0621   0.5464   1.0000
   2.750   0.6299   0.00540   0.00125  -0.0621   0.5390   1.0000
   3.000   0.6574   0.00549   0.00131  -0.0620   0.5309   1.0000
   3.250   0.6852   0.00556   0.00138  -0.0620   0.5223   1.0000
   3.500   0.7127   0.00566   0.00145  -0.0620   0.5131   1.0000
   3.750   0.7401   0.00576   0.00153  -0.0619   0.5025   1.0000
   4.000   0.7677   0.00586   0.00163  -0.0619   0.4916   1.0000
   4.250   0.7950   0.00598   0.00173  -0.0618   0.4783   1.0000
   4.500   0.8216   0.00619   0.00184  -0.0616   0.4496   1.0000
   5.000   0.8721   0.00696   0.00224  -0.0611   0.3564   1.0000
   5.250   0.8956   0.00762   0.00260  -0.0606   0.2909   1.0000
   5.500   0.9177   0.00844   0.00305  -0.0600   0.2159   1.0000
   5.750   0.9396   0.00927   0.00354  -0.0594   0.1485   1.0000
   6.000   0.9610   0.01014   0.00409  -0.0588   0.0873   1.0000
   6.250   0.9806   0.01122   0.00479  -0.0578   0.0254   1.0000
   6.500   1.0035   0.01187   0.00535  -0.0572   0.0100   1.0000
   6.750   1.0279   0.01228   0.00579  -0.0567   0.0088   1.0000
   7.000   1.0512   0.01285   0.00644  -0.0560   0.0076   1.0000
   7.250   1.0746   0.01336   0.00702  -0.0554   0.0073   1.0000
   7.500   1.0975   0.01390   0.00763  -0.0547   0.0070   1.0000
   7.750   1.1193   0.01454   0.00834  -0.0539   0.0067   1.0000
   8.000   1.1403   0.01525   0.00912  -0.0530   0.0065   1.0000
   8.250   1.1603   0.01601   0.00996  -0.0519   0.0063   1.0000
   8.500   1.1798   0.01677   0.01080  -0.0509   0.0061   1.0000
   8.750   1.1986   0.01755   0.01164  -0.0497   0.0059   1.0000
   9.000   1.2154   0.01847   0.01263  -0.0483   0.0058   1.0000
   9.250   1.2306   0.01949   0.01373  -0.0467   0.0057   1.0000
   9.500   1.2448   0.02052   0.01484  -0.0450   0.0056   1.0000
   9.750   1.2571   0.02164   0.01605  -0.0431   0.0056   1.0000
  10.000   1.2660   0.02282   0.01732  -0.0406   0.0055   1.0000
  10.250   1.2731   0.02398   0.01857  -0.0379   0.0055   1.0000
  10.500   1.2793   0.02536   0.02005  -0.0354   0.0055   1.0000
  10.750   1.2839   0.02701   0.02176  -0.0332   0.0052   1.0000
  11.000   1.2918   0.02847   0.02333  -0.0313   0.0052   1.0000
  11.250   1.2959   0.03080   0.02578  -0.0294   0.0051   1.0000
  11.500   1.2996   0.03378   0.02893  -0.0275   0.0050   1.0000
  11.750   1.3052   0.03555   0.03083  -0.0261   0.0050   1.0000
  12.000   1.3089   0.03763   0.03305  -0.0248   0.0050   1.0000
  12.250   1.3017   0.04214   0.03783  -0.0230   0.0049   1.0000
  12.500   1.2995   0.04487   0.04073  -0.0220   0.0050   1.0000
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