Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AH 83-150 Q (ah83150q-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: AH 83-150 Q (ah83150q-il)
Reynolds number: 500,000
Max Cl/Cd: 102.15 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ah83150q-il-500000-n5.txt
Download as CSV file: xf-ah83150q-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH 83-150 Q                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.1794   0.09174   0.08836  -0.0502   0.6774   0.0101
 -10.250  -0.1754   0.08837   0.08499  -0.0515   0.6759   0.0103
 -10.000  -0.1728   0.08466   0.08127  -0.0529   0.6745   0.0104
  -9.750  -0.1710   0.08085   0.07746  -0.0544   0.6732   0.0108
  -9.500  -0.1724   0.07624   0.07285  -0.0563   0.6721   0.0108
  -9.250  -0.2445   0.07566   0.07211  -0.0677   0.6771   0.0098
  -8.250  -0.2901   0.05507   0.05131  -0.0779   0.6704   0.0098
  -7.750  -0.2946   0.05086   0.04701  -0.0767   0.6671   0.0094
  -7.500  -0.2925   0.04808   0.04413  -0.0762   0.6652   0.0090
  -7.250  -0.2901   0.04398   0.03984  -0.0754   0.6634   0.0089
  -7.000  -0.2838   0.04020   0.03583  -0.0744   0.6615   0.0085
  -6.750  -0.2748   0.03630   0.03166  -0.0731   0.6598   0.0084
  -6.500  -0.2609   0.03320   0.02829  -0.0720   0.6582   0.0088
  -6.250  -0.2454   0.02995   0.02471  -0.0707   0.6567   0.0091
  -6.000  -0.2284   0.02708   0.02152  -0.0696   0.6552   0.0087
  -5.750  -0.2123   0.02154   0.01510  -0.0670   0.6538   0.0074
  -5.250  -0.1621   0.01919   0.01234  -0.0663   0.6502   0.0072
  -5.000  -0.1367   0.01796   0.01094  -0.0660   0.6484   0.0072
  -4.750  -0.1106   0.01691   0.00973  -0.0658   0.6467   0.0072
  -4.500  -0.0842   0.01602   0.00871  -0.0656   0.6448   0.0072
  -4.250  -0.0578   0.01527   0.00785  -0.0654   0.6430   0.0072
  -4.000  -0.0316   0.01454   0.00704  -0.0651   0.6413   0.0072
  -3.750  -0.0059   0.01381   0.00625  -0.0649   0.6398   0.0074
  -3.500   0.0201   0.01326   0.00565  -0.0647   0.6382   0.0075
  -3.250   0.0466   0.01280   0.00519  -0.0646   0.6364   0.0077
  -3.000   0.0735   0.01242   0.00480  -0.0646   0.6344   0.0080
  -2.750   0.1006   0.01207   0.00444  -0.0646   0.6325   0.0085
  -2.500   0.1280   0.01178   0.00412  -0.0647   0.6306   0.0090
  -2.250   0.1557   0.01154   0.00384  -0.0647   0.6288   0.0097
  -2.000   0.1833   0.01128   0.00355  -0.0648   0.6269   0.0104
  -1.750   0.2112   0.01111   0.00334  -0.0650   0.6251   0.0119
  -1.500   0.2392   0.01096   0.00315  -0.0652   0.6234   0.0144
  -1.250   0.2674   0.01078   0.00300  -0.0653   0.6215   0.0202
  -1.000   0.2933   0.01015   0.00285  -0.0655   0.6193   0.1554
  -0.750   0.3094   0.00807   0.00276  -0.0647   0.6169   0.7128
  -0.500   0.3346   0.00848   0.00332  -0.0635   0.6145   0.8011
  -0.250   0.3597   0.00890   0.00373  -0.0624   0.6121   0.8211
   0.000   0.3871   0.00911   0.00388  -0.0621   0.6099   0.8316
   0.250   0.4146   0.00919   0.00392  -0.0619   0.6077   0.8346
   0.500   0.4432   0.00922   0.00393  -0.0622   0.6051   0.8365
   0.750   0.4724   0.00922   0.00390  -0.0626   0.6022   0.8376
   1.000   0.5016   0.00922   0.00387  -0.0630   0.5994   0.8388
   1.250   0.5308   0.00923   0.00383  -0.0635   0.5967   0.8401
   1.500   0.5600   0.00925   0.00379  -0.0639   0.5941   0.8411
   1.750   0.5892   0.00926   0.00379  -0.0644   0.5911   0.8421
   2.000   0.6185   0.00928   0.00381  -0.0649   0.5876   0.8434
   2.250   0.6477   0.00930   0.00380  -0.0655   0.5839   0.8445
   2.500   0.6769   0.00932   0.00379  -0.0660   0.5805   0.8454
   2.750   0.7055   0.00934   0.00380  -0.0663   0.5770   0.8460
   3.000   0.7341   0.00936   0.00384  -0.0667   0.5728   0.8466
   3.250   0.7624   0.00938   0.00388  -0.0670   0.5686   0.8474
   3.500   0.7905   0.00942   0.00389  -0.0673   0.5646   0.8482
   3.750   0.8188   0.00946   0.00396  -0.0676   0.5600   0.8490
   4.000   0.8470   0.00951   0.00402  -0.0679   0.5549   0.8498
   4.250   0.8750   0.00956   0.00407  -0.0682   0.5500   0.8505
   4.500   0.9031   0.00962   0.00416  -0.0685   0.5443   0.8512
   4.750   0.9308   0.00969   0.00423  -0.0688   0.5377   0.8520
   5.000   0.9583   0.00976   0.00432  -0.0690   0.5305   0.8529
   5.250   0.9848   0.00987   0.00440  -0.0691   0.5194   0.8539
   5.500   1.0107   0.01001   0.00451  -0.0691   0.5039   0.8550
   5.750   1.0360   0.01020   0.00465  -0.0690   0.4870   0.8559
   6.000   1.0612   0.01040   0.00483  -0.0689   0.4739   0.8568
   6.250   1.0859   0.01063   0.00503  -0.0687   0.4601   0.8576
   6.500   1.1096   0.01091   0.00527  -0.0685   0.4437   0.8584
   6.750   1.1326   0.01122   0.00554  -0.0681   0.4274   0.8593
   7.250   1.1745   0.01194   0.00619  -0.0666   0.3939   0.8611
   7.500   1.1929   0.01237   0.00658  -0.0655   0.3761   0.8621
   7.750   1.2105   0.01281   0.00699  -0.0642   0.3600   0.8631
   8.000   1.2260   0.01330   0.00745  -0.0627   0.3436   0.8641
   8.500   1.2413   0.01454   0.00859  -0.0570   0.3079   0.8668
   8.750   1.2432   0.01549   0.00947  -0.0538   0.2866   0.8685
   9.000   1.2419   0.01678   0.01065  -0.0506   0.2625   0.8702
   9.250   1.2418   0.01822   0.01198  -0.0480   0.2389   0.8721
   9.500   1.2447   0.01963   0.01332  -0.0459   0.2192   0.8737
   9.750   1.2483   0.02106   0.01469  -0.0441   0.2006   0.8750
  10.000   1.2529   0.02247   0.01606  -0.0424   0.1844   0.8762
  10.250   1.2586   0.02385   0.01741  -0.0409   0.1706   0.8775
  10.500   1.2645   0.02526   0.01879  -0.0395   0.1581   0.8787
  10.750   1.2712   0.02665   0.02017  -0.0382   0.1479   0.8800
  11.000   1.2778   0.02809   0.02160  -0.0370   0.1379   0.8812
  11.250   1.2863   0.02941   0.02294  -0.0361   0.1296   0.8825
  11.500   1.2930   0.03090   0.02442  -0.0350   0.1214   0.8838
  11.750   1.3003   0.03238   0.02591  -0.0341   0.1132   0.8851
  12.000   1.3085   0.03382   0.02737  -0.0333   0.1063   0.8865
  12.250   1.3133   0.03557   0.02910  -0.0324   0.0967   0.8879
  12.500   1.3232   0.03696   0.03053  -0.0318   0.0912   0.8891
  12.750   1.3286   0.03876   0.03234  -0.0310   0.0824   0.8904
  13.000   1.3334   0.04065   0.03420  -0.0303   0.0729   0.8917
  13.250   1.3391   0.04251   0.03608  -0.0298   0.0650   0.8930
  13.500   1.3447   0.04443   0.03801  -0.0293   0.0580   0.8944
  13.750   1.3490   0.04654   0.04012  -0.0288   0.0518   0.8959
  14.000   1.3551   0.04852   0.04214  -0.0285   0.0472   0.8974
  14.250   1.3590   0.05076   0.04441  -0.0282   0.0426   0.8990
  14.500   1.3653   0.05282   0.04653  -0.0280   0.0392   0.9006
  14.750   1.3701   0.05510   0.04886  -0.0279   0.0368   0.9022
  15.250   1.3814   0.05959   0.05348  -0.0279   0.0321   0.9062
  15.500   1.3860   0.06203   0.05600  -0.0281   0.0302   0.9085
  15.750   1.3904   0.06460   0.05863  -0.0284   0.0288   0.9111
  16.000   1.3954   0.06712   0.06125  -0.0287   0.0273   0.9141
  16.500   1.4025   0.07265   0.06694  -0.0297   0.0234   0.9233
  17.000   1.4113   0.07854   0.07306  -0.0317   0.0205   0.9648
  17.250   1.4116   0.08156   0.07616  -0.0320   0.0189   1.0000
  17.500   1.4120   0.08492   0.07957  -0.0329   0.0170   1.0000
  17.750   1.4129   0.08826   0.08300  -0.0338   0.0152   1.0000
  18.000   1.4116   0.09196   0.08676  -0.0348   0.0132   1.0000
  18.250   1.4085   0.09598   0.09082  -0.0360   0.0102   1.0000
  18.500   1.4029   0.10044   0.09534  -0.0375   0.0076   1.0000
  18.750   1.3961   0.10518   0.10015  -0.0391   0.0060   1.0000
  19.000   1.3928   0.10945   0.10451  -0.0407   0.0055   1.0000
<< Back to AH 83-150 Q (ah83150q-il)

Polar data table (+)

Polar graphs


<< Back to AH 83-150 Q (ah83150q-il)