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AH21 9% version (Andrew Hollom) (ah21-9-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AH21 9% version (Andrew Hollom) (ah21-9-il)
Reynolds number: 500,000
Max Cl/Cd: 103.18 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ah21-9-il-500000.txt
Download as CSV file: xf-ah21-9-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH21 9% version (Andrew Hollom)                 
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.5610   0.06498   0.06282  -0.0334   1.0000   0.0154
  -7.250  -0.5662   0.06069   0.05847  -0.0337   1.0000   0.0156
  -7.000  -0.5640   0.05899   0.05675  -0.0325   1.0000   0.0162
  -6.750  -0.5611   0.05506   0.05273  -0.0327   1.0000   0.0164
  -6.500  -0.5548   0.05206   0.04964  -0.0324   1.0000   0.0170
  -6.250  -0.5469   0.04824   0.04568  -0.0324   1.0000   0.0175
  -6.000  -0.5271   0.04345   0.04067  -0.0346   0.9988   0.0184
  -5.750  -0.4988   0.03842   0.03531  -0.0375   0.9966   0.0201
  -5.500  -0.4622   0.03662   0.03309  -0.0388   0.9948   0.0223
  -5.250  -0.4433   0.02794   0.02372  -0.0410   0.9924   0.0235
  -5.000  -0.4171   0.02575   0.02144  -0.0421   0.9903   0.0253
  -4.750  -0.3870   0.02138   0.01645  -0.0419   0.9888   0.0206
  -4.500  -0.3563   0.01933   0.01405  -0.0425   0.9872   0.0213
  -4.250  -0.3254   0.01713   0.01150  -0.0430   0.9860   0.0213
  -4.000  -0.2938   0.01561   0.00974  -0.0437   0.9850   0.0217
  -3.750  -0.2602   0.01487   0.00886  -0.0449   0.9838   0.0223
  -3.500  -0.2298   0.01304   0.00686  -0.0455   0.9832   0.0240
  -3.250  -0.2030   0.01219   0.00595  -0.0454   0.9803   0.0253
  -3.000  -0.1731   0.01163   0.00535  -0.0459   0.9775   0.0267
  -2.750  -0.1408   0.01121   0.00488  -0.0469   0.9752   0.0283
  -2.500  -0.1068   0.01090   0.00453  -0.0483   0.9732   0.0310
  -2.250  -0.0715   0.01056   0.00412  -0.0499   0.9717   0.0342
  -2.000  -0.0356   0.01011   0.00381  -0.0517   0.9704   0.0726
  -1.750  -0.0112   0.00908   0.00365  -0.0517   0.9663   0.3090
  -1.500   0.0172   0.00823   0.00359  -0.0524   0.9629   0.5407
  -1.250   0.0469   0.00759   0.00364  -0.0528   0.9603   0.7316
  -1.000   0.0791   0.00729   0.00365  -0.0533   0.9582   0.8308
  -0.750   0.1124   0.00711   0.00365  -0.0539   0.9565   0.9067
  -0.500   0.1558   0.00703   0.00365  -0.0567   0.9548   0.9757
  -0.250   0.2068   0.00693   0.00353  -0.0617   0.9534   0.9925
   0.000   0.2498   0.00679   0.00336  -0.0649   0.9497   1.0000
   0.250   0.2942   0.00656   0.00312  -0.0682   0.9468   1.0000
   0.500   0.3435   0.00629   0.00285  -0.0726   0.9444   1.0000
   0.750   0.3739   0.00606   0.00261  -0.0729   0.9359   1.0000
   1.000   0.4160   0.00577   0.00233  -0.0757   0.9296   1.0000
   1.250   0.4427   0.00558   0.00215  -0.0751   0.9168   1.0000
   1.500   0.4720   0.00542   0.00198  -0.0752   0.9014   1.0000
   1.750   0.4996   0.00531   0.00185  -0.0748   0.8770   1.0000
   2.000   0.5287   0.00527   0.00175  -0.0748   0.8466   1.0000
   2.250   0.5541   0.00537   0.00168  -0.0739   0.7948   1.0000
   2.500   0.5730   0.00567   0.00169  -0.0717   0.7246   1.0000
   2.750   0.5857   0.00623   0.00180  -0.0683   0.6259   1.0000
   3.000   0.5973   0.00699   0.00203  -0.0649   0.5036   1.0000
   3.250   0.6120   0.00781   0.00230  -0.0625   0.3789   1.0000
   3.500   0.6295   0.00863   0.00261  -0.0607   0.2636   1.0000
   3.750   0.6497   0.00933   0.00292  -0.0595   0.1783   1.0000
   4.000   0.6713   0.00994   0.00323  -0.0585   0.1186   1.0000
   4.250   0.6940   0.01046   0.00359  -0.0576   0.0824   1.0000
   4.500   0.7175   0.01091   0.00396  -0.0568   0.0616   1.0000
   4.750   0.7409   0.01141   0.00440  -0.0560   0.0478   1.0000
   5.000   0.7643   0.01191   0.00492  -0.0551   0.0392   1.0000
   5.250   0.7852   0.01278   0.00581  -0.0538   0.0323   1.0000
   5.500   0.8094   0.01318   0.00629  -0.0531   0.0304   1.0000
   5.750   0.8326   0.01376   0.00693  -0.0522   0.0282   1.0000
   6.000   0.8559   0.01430   0.00751  -0.0514   0.0258   1.0000
   6.250   0.8752   0.01562   0.00891  -0.0498   0.0232   1.0000
   6.500   0.8970   0.01677   0.01018  -0.0487   0.0221   1.0000
   6.750   0.9210   0.01737   0.01089  -0.0480   0.0208   1.0000
   7.000   0.9445   0.01796   0.01156  -0.0473   0.0190   1.0000
   7.250   0.9674   0.01870   0.01237  -0.0465   0.0176   1.0000
   7.500   0.9883   0.02009   0.01386  -0.0454   0.0162   1.0000
   7.750   1.0062   0.02313   0.01726  -0.0438   0.0149   1.0000
   8.000   1.0286   0.02369   0.01797  -0.0429   0.0139   1.0000
   8.250   1.0493   0.02469   0.01916  -0.0418   0.0127   1.0000
   8.500   1.0684   0.02592   0.02056  -0.0405   0.0118   1.0000
   8.750   1.0863   0.02720   0.02199  -0.0392   0.0112   1.0000
   9.000   1.0994   0.02944   0.02445  -0.0373   0.0107   1.0000
   9.250   1.0915   0.03562   0.03129  -0.0328   0.0102   1.0000
   9.500   1.0812   0.04086   0.03706  -0.0284   0.0101   1.0000
   9.750   1.0796   0.04386   0.04041  -0.0252   0.0099   1.0000
  10.000   1.0641   0.04813   0.04501  -0.0209   0.0099   1.0000
  10.250   1.0384   0.05224   0.04937  -0.0159   0.0100   1.0000
  10.500   0.9714   0.06555   0.06327  -0.0120   0.0118   1.0000
  10.750   0.9555   0.06952   0.06734  -0.0127   0.0115   1.0000
  11.000   0.9367   0.07530   0.07323  -0.0160   0.0114   1.0000
  11.250   0.9384   0.07737   0.07532  -0.0184   0.0111   1.0000
  11.500   0.9354   0.08236   0.08037  -0.0236   0.0108   1.0000
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