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

AH 85-L-120/17 (ah85l120-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AH 85-L-120/17 (ah85l120-il)
Reynolds number: 500,000
Max Cl/Cd: 48.33 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ah85l120-il-500000.txt
Download as CSV file: xf-ah85l120-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH 85-L-120/17                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.7020   0.02394   0.01822  -0.0433   0.9403   0.0147
  -8.250  -0.6738   0.02111   0.01509  -0.0435   0.9382   0.0142
  -8.000  -0.6556   0.02110   0.01506  -0.0425   0.9342   0.0156
  -7.750  -0.6278   0.01951   0.01329  -0.0428   0.9320   0.0163
  -7.500  -0.5995   0.01810   0.01175  -0.0432   0.9301   0.0164
  -7.250  -0.5742   0.01698   0.01053  -0.0431   0.9280   0.0169
  -7.000  -0.5544   0.01631   0.00978  -0.0419   0.9252   0.0179
  -6.750  -0.5353   0.01573   0.00913  -0.0405   0.9220   0.0184
  -6.500  -0.5161   0.01521   0.00855  -0.0391   0.9191   0.0191
  -6.250  -0.5072   0.01398   0.00718  -0.0358   0.9159   0.0210
  -6.000  -0.4904   0.01356   0.00675  -0.0340   0.9131   0.0231
  -5.750  -0.4732   0.01317   0.00630  -0.0322   0.9101   0.0252
  -5.500  -0.4546   0.01280   0.00588  -0.0305   0.9073   0.0281
  -5.250  -0.4381   0.01227   0.00532  -0.0285   0.9045   0.0382
  -5.000  -0.4243   0.01160   0.00490  -0.0261   0.9018   0.0872
  -4.750  -0.4153   0.01080   0.00456  -0.0230   0.8987   0.1888
  -4.500  -0.4083   0.00993   0.00424  -0.0195   0.8950   0.3110
  -4.250  -0.4060   0.00896   0.00393  -0.0150   0.8913   0.4609
  -4.000  -0.3995   0.00829   0.00374  -0.0108   0.8883   0.5797
  -3.750  -0.3858   0.00796   0.00375  -0.0079   0.8856   0.6615
  -3.500  -0.3654   0.00790   0.00383  -0.0063   0.8827   0.7083
  -3.250  -0.3431   0.00800   0.00397  -0.0049   0.8799   0.7462
  -3.000  -0.3193   0.00822   0.00420  -0.0038   0.8774   0.7716
  -2.750  -0.2927   0.00838   0.00433  -0.0034   0.8753   0.7826
  -2.500  -0.2663   0.00857   0.00448  -0.0029   0.8733   0.7921
  -2.250  -0.2421   0.00871   0.00457  -0.0022   0.8703   0.8005
  -2.000  -0.2148   0.00885   0.00471  -0.0020   0.8673   0.8064
  -1.750  -0.1889   0.00901   0.00484  -0.0014   0.8644   0.8143
  -1.500  -0.1607   0.00915   0.00497  -0.0013   0.8621   0.8199
  -1.250  -0.1325   0.00933   0.00512  -0.0012   0.8602   0.8260
  -1.000  -0.1077   0.00943   0.00520  -0.0006   0.8575   0.8318
  -0.750  -0.0805   0.00949   0.00528  -0.0005   0.8541   0.8347
  -0.500  -0.0533   0.00953   0.00532  -0.0004   0.8509   0.8382
  -0.250  -0.0269   0.00953   0.00529  -0.0002   0.8482   0.8424
   0.000   0.0000   0.00950   0.00523   0.0000   0.8458   0.8458
   0.250   0.0269   0.00953   0.00528   0.0002   0.8424   0.8481
   0.500   0.0533   0.00953   0.00532   0.0004   0.8381   0.8509
   0.750   0.0806   0.00949   0.00528   0.0005   0.8347   0.8541
   1.000   0.1078   0.00942   0.00518   0.0006   0.8317   0.8575
   1.250   0.1326   0.00933   0.00513   0.0012   0.8261   0.8602
   1.500   0.1608   0.00915   0.00497   0.0013   0.8198   0.8622
   1.750   0.1890   0.00902   0.00485   0.0014   0.8143   0.8645
   2.000   0.2148   0.00887   0.00473   0.0019   0.8068   0.8673
   2.250   0.2421   0.00871   0.00457   0.0021   0.8006   0.8703
   2.500   0.2664   0.00857   0.00448   0.0029   0.7922   0.8734
   2.750   0.2928   0.00838   0.00433   0.0034   0.7826   0.8753
   3.000   0.3197   0.00822   0.00421   0.0037   0.7726   0.8774
   3.250   0.3432   0.00801   0.00401   0.0049   0.7486   0.8798
   3.500   0.3658   0.00790   0.00386   0.0062   0.7123   0.8826
   3.750   0.3852   0.00797   0.00374   0.0081   0.6568   0.8856
   4.000   0.3978   0.00834   0.00375   0.0112   0.5696   0.8884
   4.250   0.4056   0.00898   0.00393   0.0151   0.4579   0.8914
   4.500   0.4102   0.00985   0.00423   0.0192   0.3233   0.8949
   4.750   0.4170   0.01070   0.00455   0.0227   0.2023   0.8986
   5.000   0.4262   0.01150   0.00487   0.0258   0.1006   0.9017
   5.250   0.4378   0.01229   0.00534   0.0286   0.0365   0.9046
   5.500   0.4554   0.01275   0.00582   0.0304   0.0288   0.9073
   5.750   0.4731   0.01317   0.00631   0.0322   0.0256   0.9101
   6.000   0.4916   0.01348   0.00664   0.0338   0.0223   0.9131
   6.250   0.5043   0.01417   0.00738   0.0364   0.0199   0.9161
   6.500   0.5173   0.01507   0.00840   0.0389   0.0188   0.9190
   6.750   0.5363   0.01553   0.00893   0.0403   0.0183   0.9219
   7.000   0.5547   0.01609   0.00956   0.0419   0.0170   0.9253
   7.250   0.5744   0.01706   0.01061   0.0431   0.0171   0.9280
   7.500   0.5993   0.01809   0.01174   0.0433   0.0163   0.9302
   7.750   0.6277   0.01951   0.01329   0.0428   0.0162   0.9320
   8.000   0.6574   0.02159   0.01560   0.0423   0.0158   0.9339
   8.250   0.6752   0.02153   0.01554   0.0434   0.0142   0.9379
   8.500   0.6988   0.02309   0.01727   0.0436   0.0140   0.9408
   8.750   0.7217   0.02629   0.02087   0.0443   0.0145   0.9434
<< Back to AH 85-L-120/17 (ah85l120-il)

Polar data table (+)

Polar graphs


<< Back to AH 85-L-120/17 (ah85l120-il)