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AG16 (ag16-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: AG16 (ag16-il)
Reynolds number: 1,000,000
Max Cl/Cd: 95.34 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag16-il-1000000.txt
Download as CSV file: xf-ag16-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG16                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5694   0.10135   0.09979   0.0146   1.0000   0.0093
  -8.750  -0.5660   0.09711   0.09557   0.0125   1.0000   0.0093
  -8.500  -0.5631   0.09275   0.09122   0.0101   1.0000   0.0094
  -6.500  -0.4975   0.01975   0.01615  -0.0495   1.0000   0.0059
  -6.250  -0.4715   0.01845   0.01466  -0.0494   1.0000   0.0057
  -6.000  -0.4457   0.01698   0.01299  -0.0494   1.0000   0.0056
  -5.750  -0.4202   0.01460   0.01028  -0.0493   1.0000   0.0056
  -5.500  -0.3953   0.01111   0.00633  -0.0492   1.0000   0.0063
  -5.250  -0.3690   0.01055   0.00573  -0.0490   1.0000   0.0071
  -5.000  -0.3426   0.01003   0.00513  -0.0487   1.0000   0.0077
  -4.750  -0.3163   0.00950   0.00453  -0.0483   1.0000   0.0082
  -4.500  -0.2900   0.00900   0.00396  -0.0479   1.0000   0.0088
  -4.250  -0.2639   0.00867   0.00359  -0.0475   1.0000   0.0093
  -4.000  -0.2372   0.00794   0.00273  -0.0472   1.0000   0.0116
  -3.750  -0.2074   0.00760   0.00236  -0.0475   0.9989   0.0148
  -3.500  -0.1710   0.00711   0.00189  -0.0493   0.9956   0.0309
  -3.250  -0.1355   0.00675   0.00165  -0.0510   0.9913   0.0612
  -3.000  -0.1006   0.00641   0.00146  -0.0526   0.9844   0.1005
  -2.750  -0.0667   0.00608   0.00130  -0.0540   0.9749   0.1486
  -2.500  -0.0343   0.00580   0.00117  -0.0549   0.9611   0.1971
  -2.250  -0.0055   0.00559   0.00106  -0.0550   0.9419   0.2435
  -2.000   0.0209   0.00545   0.00098  -0.0544   0.9185   0.2805
  -1.750   0.0468   0.00536   0.00091  -0.0538   0.8928   0.3162
  -1.500   0.0731   0.00529   0.00084  -0.0532   0.8653   0.3529
  -1.250   0.0998   0.00527   0.00079  -0.0528   0.8366   0.3847
  -1.000   0.1267   0.00524   0.00075  -0.0524   0.8070   0.4208
  -0.750   0.1538   0.00520   0.00073  -0.0522   0.7772   0.4678
  -0.500   0.1811   0.00519   0.00071  -0.0519   0.7479   0.5098
   0.000   0.2360   0.00514   0.00072  -0.0516   0.6895   0.6058
   0.250   0.2633   0.00509   0.00074  -0.0514   0.6612   0.6651
   0.500   0.2901   0.00497   0.00078  -0.0511   0.6328   0.7473
   0.750   0.3118   0.00463   0.00081  -0.0494   0.6059   0.8991
   1.000   0.3418   0.00463   0.00080  -0.0497   0.5773   1.0000
   1.250   0.3697   0.00478   0.00083  -0.0496   0.5496   1.0000
   1.500   0.3976   0.00494   0.00088  -0.0495   0.5214   1.0000
   1.750   0.4255   0.00510   0.00093  -0.0495   0.4933   1.0000
   2.000   0.4533   0.00528   0.00099  -0.0494   0.4653   1.0000
   2.250   0.4811   0.00546   0.00106  -0.0494   0.4373   1.0000
   2.500   0.5089   0.00565   0.00114  -0.0493   0.4090   1.0000
   2.750   0.5366   0.00586   0.00125  -0.0492   0.3800   1.0000
   3.000   0.5642   0.00608   0.00136  -0.0492   0.3504   1.0000
   3.250   0.5917   0.00632   0.00148  -0.0491   0.3206   1.0000
   3.500   0.6192   0.00655   0.00161  -0.0491   0.2941   1.0000
   3.750   0.6466   0.00681   0.00178  -0.0490   0.2647   1.0000
   4.000   0.6739   0.00708   0.00194  -0.0489   0.2360   1.0000
   4.250   0.7012   0.00736   0.00212  -0.0488   0.2098   1.0000
   4.500   0.7284   0.00764   0.00232  -0.0487   0.1869   1.0000
   4.750   0.7554   0.00797   0.00255  -0.0486   0.1594   1.0000
   5.000   0.7822   0.00832   0.00279  -0.0485   0.1335   1.0000
   5.250   0.8092   0.00864   0.00304  -0.0484   0.1129   1.0000
   5.500   0.8358   0.00903   0.00333  -0.0483   0.0911   1.0000
   5.750   0.8621   0.00949   0.00366  -0.0481   0.0670   1.0000
   6.000   0.8884   0.00992   0.00401  -0.0479   0.0486   1.0000
   6.250   0.9146   0.01036   0.00440  -0.0477   0.0352   1.0000
   6.500   0.9409   0.01078   0.00479  -0.0475   0.0259   1.0000
   6.750   0.9671   0.01122   0.00522  -0.0473   0.0186   1.0000
   7.000   0.9928   0.01176   0.00576  -0.0469   0.0123   1.0000
   7.250   1.0177   0.01256   0.00659  -0.0464   0.0076   1.0000
   7.500   1.0433   0.01308   0.00718  -0.0461   0.0063   1.0000
   7.750   1.0681   0.01382   0.00797  -0.0457   0.0052   1.0000
   8.000   1.0904   0.01513   0.00950  -0.0449   0.0045   1.0000
   8.250   1.1149   0.01580   0.01027  -0.0444   0.0044   1.0000
   8.500   1.1384   0.01665   0.01123  -0.0438   0.0042   1.0000
   8.750   1.1613   0.01760   0.01230  -0.0432   0.0040   1.0000
   9.000   1.1835   0.01863   0.01346  -0.0425   0.0039   1.0000
   9.250   1.2052   0.01972   0.01469  -0.0418   0.0037   1.0000
   9.500   1.2262   0.02085   0.01595  -0.0410   0.0035   1.0000
   9.750   1.2470   0.02197   0.01719  -0.0403   0.0033   1.0000
  10.000   1.2674   0.02305   0.01839  -0.0395   0.0031   1.0000
  10.250   1.2868   0.02421   0.01968  -0.0388   0.0029   1.0000
  10.500   1.3017   0.02605   0.02170  -0.0376   0.0027   1.0000
  10.750   1.2971   0.03090   0.02702  -0.0348   0.0025   1.0000
  11.250   1.2770   0.03956   0.03644  -0.0297   0.0024   1.0000
  11.500   1.2677   0.04245   0.03953  -0.0279   0.0024   1.0000
  11.750   1.2550   0.04643   0.04370  -0.0281   0.0024   1.0000
  12.000   1.2394   0.05203   0.04951  -0.0311   0.0024   1.0000
  12.250   1.2179   0.06016   0.05786  -0.0371   0.0024   1.0000
  12.500   1.1950   0.06970   0.06758  -0.0442   0.0024   1.0000
  12.750   1.1844   0.07704   0.07503  -0.0494   0.0024   1.0000
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