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AH 93-156 (ah93156-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AH 93-156 (ah93156-il)
Reynolds number: 50,000
Max Cl/Cd: 4.8 at α=9.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ah93156-il-50000.txt
Download as CSV file: xf-ah93156-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH 93-156                                       
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4443   0.12803   0.12251  -0.0174   1.0000   0.2678
  -8.750  -0.4358   0.12477   0.11926  -0.0154   1.0000   0.2778
  -8.500  -0.4723   0.12443   0.11905  -0.0142   1.0000   0.2852
  -8.250  -0.4572   0.12091   0.11553  -0.0119   1.0000   0.2990
  -8.000  -0.4558   0.11807   0.11270  -0.0099   1.0000   0.3095
  -7.750  -0.5037   0.11803   0.11283  -0.0077   1.0000   0.3180
  -7.500  -0.4949   0.11492   0.10971  -0.0053   1.0000   0.3334
  -7.250  -0.4888   0.11195   0.10675  -0.0030   1.0000   0.3482
  -7.000  -0.4881   0.10931   0.10414  -0.0006   1.0000   0.3626
  -6.750  -0.4993   0.10728   0.10217   0.0025   1.0000   0.3801
  -6.500  -0.5163   0.10557   0.10053   0.0061   1.0000   0.3980
  -6.250  -0.4961   0.10264   0.09758   0.0089   1.0000   0.4268
  -6.000  -0.4940   0.10038   0.09536   0.0123   1.0000   0.4543
  -5.750  -0.4934   0.09837   0.09338   0.0162   1.0000   0.4852
  -5.000  -0.6493   0.06657   0.05990  -0.0126   1.0000   0.1719
  -4.750  -0.6354   0.06121   0.05401  -0.0125   1.0000   0.1545
  -4.500  -0.6216   0.05760   0.05005  -0.0118   1.0000   0.1510
  -4.250  -0.6058   0.05413   0.04607  -0.0111   1.0000   0.1480
  -4.000  -0.5872   0.05085   0.04207  -0.0103   1.0000   0.1446
  -3.750  -0.5672   0.04837   0.03890  -0.0094   1.0000   0.1431
  -3.500  -0.5494   0.04659   0.03706  -0.0086   1.0000   0.1478
  -3.250  -0.5295   0.04496   0.03506  -0.0077   1.0000   0.1521
  -3.000  -0.5080   0.04343   0.03303  -0.0068   1.0000   0.1557
  -2.750  -0.4879   0.04207   0.03155  -0.0061   1.0000   0.1628
  -2.500  -0.4667   0.04117   0.03031  -0.0053   1.0000   0.1720
  -2.250  -0.4458   0.04018   0.02931  -0.0045   1.0000   0.1820
  -2.000  -0.4245   0.03950   0.02859  -0.0037   1.0000   0.1975
  -1.750  -0.4018   0.03902   0.02810  -0.0028   1.0000   0.2174
  -1.500  -0.2315   0.04099   0.03330  -0.0203   1.0000   1.0000
  -1.250  -0.2225   0.04116   0.03311  -0.0183   1.0000   1.0000
  -1.000  -0.2137   0.04136   0.03299  -0.0164   1.0000   1.0000
  -0.750  -0.2050   0.04159   0.03295  -0.0144   1.0000   1.0000
  -0.500  -0.1963   0.04183   0.03297  -0.0125   1.0000   1.0000
  -0.250  -0.1874   0.04211   0.03303  -0.0107   1.0000   1.0000
   0.000  -0.1786   0.04242   0.03313  -0.0088   1.0000   1.0000
   0.250  -0.1696   0.04276   0.03329  -0.0071   1.0000   1.0000
   0.500  -0.1605   0.04313   0.03349  -0.0054   1.0000   1.0000
   0.750  -0.1512   0.04353   0.03373  -0.0037   1.0000   1.0000
   1.000  -0.1418   0.04397   0.03403  -0.0022   1.0000   1.0000
   1.250  -0.1322   0.04445   0.03435  -0.0007   1.0000   1.0000
   1.500  -0.1224   0.04498   0.03476   0.0007   1.0000   1.0000
   1.750  -0.1121   0.04556   0.03521   0.0020   1.0000   1.0000
   2.000  -0.1014   0.04621   0.03575   0.0031   1.0000   1.0000
   2.250  -0.0805   0.04754   0.03694   0.0021   0.9962   1.0000
   2.500  -0.0447   0.05001   0.03925  -0.0017   0.9847   1.0000
   2.750  -0.0117   0.05226   0.04137  -0.0050   0.9717   1.0000
   3.000   0.0178   0.05422   0.04322  -0.0075   0.9578   1.0000
   3.250   0.0445   0.05596   0.04488  -0.0094   0.9434   1.0000
   3.500   0.0681   0.05751   0.04635  -0.0107   0.9293   1.0000
   3.750   0.0898   0.05900   0.04778  -0.0117   0.9154   1.0000
   4.000   0.1098   0.06049   0.04922  -0.0123   0.9027   1.0000
   4.250   0.1309   0.06219   0.05088  -0.0132   0.8909   1.0000
   4.500   0.1613   0.06488   0.05350  -0.0155   0.8803   1.0000
   4.750   0.1803   0.06621   0.05481  -0.0159   0.8668   1.0000
   5.000   0.1935   0.06722   0.05581  -0.0155   0.8540   1.0000
   5.250   0.2082   0.06865   0.05723  -0.0154   0.8431   1.0000
   5.500   0.2350   0.07122   0.05978  -0.0171   0.8337   1.0000
   5.750   0.2552   0.07289   0.06145  -0.0178   0.8209   1.0000
   6.000   0.2635   0.07384   0.06242  -0.0168   0.8093   1.0000
   6.250   0.2808   0.07580   0.06440  -0.0173   0.7999   1.0000
   6.500   0.3116   0.07867   0.06729  -0.0195   0.7889   1.0000
   6.750   0.3154   0.07933   0.06798  -0.0181   0.7770   1.0000
   7.000   0.3297   0.08128   0.06996  -0.0183   0.7683   1.0000
   7.250   0.3638   0.08460   0.07333  -0.0210   0.7572   1.0000
   7.500   0.3635   0.08509   0.07386  -0.0193   0.7455   1.0000
   7.750   0.3756   0.08706   0.07588  -0.0194   0.7366   1.0000
   8.000   0.4078   0.09037   0.07925  -0.0217   0.7258   1.0000
   8.250   0.4095   0.09123   0.08017  -0.0205   0.7142   1.0000
   8.500   0.4185   0.09320   0.08222  -0.0205   0.7050   1.0000
   8.750   0.4472   0.09643   0.08552  -0.0224   0.6950   1.0000
   9.000   0.4546   0.09787   0.08704  -0.0220   0.6828   1.0000
   9.250   0.4592   0.09976   0.08899  -0.0217   0.6735   1.0000
   9.500   0.4821   0.10282   0.09216  -0.0231   0.6641   1.0000
   9.750   0.5066   0.10560   0.09505  -0.0244   0.6508   1.0000
  10.000   0.4982   0.10676   0.09627  -0.0232   0.6418   1.0000
  10.250   0.5134   0.10958   0.09918  -0.0240   0.6327   1.0000
  10.500   0.5374   0.11264   0.10235  -0.0253   0.6204   1.0000
  10.750   0.5414   0.11442   0.10424  -0.0252   0.6087   1.0000
  11.000   0.5451   0.11679   0.10670  -0.0253   0.5994   1.0000
  11.250   0.5658   0.12016   0.11018  -0.0266   0.5899   1.0000
  11.500   0.5859   0.12306   0.11320  -0.0275   0.5760   1.0000
  11.750   0.5820   0.12479   0.11501  -0.0272   0.5655   1.0000
  12.000   0.5892   0.12751   0.11785  -0.0277   0.5550   1.0000
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