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PT40 (pt40-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: PT40 (pt40-il)
Reynolds number: 50,000
Max Cl/Cd: 32.06 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pt40-il-50000.txt
Download as CSV file: xf-pt40-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: PT40                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3309   0.10682   0.09914  -0.0181   1.0000   0.2496
  -7.250  -0.3000   0.10223   0.09455  -0.0161   1.0000   0.2602
  -7.000  -0.3366   0.10360   0.09613  -0.0156   1.0000   0.2686
  -6.750  -0.2997   0.09763   0.09016  -0.0142   1.0000   0.2778
  -6.500  -0.3173   0.09725   0.08994  -0.0124   1.0000   0.2890
  -6.250  -0.3126   0.09401   0.08680  -0.0109   1.0000   0.2966
  -6.000  -0.3114   0.09224   0.08514  -0.0085   1.0000   0.3090
  -5.750  -0.3576   0.09284   0.08599  -0.0060   1.0000   0.3152
  -5.500  -0.3216   0.08853   0.08169  -0.0030   1.0000   0.3304
  -5.250  -0.3477   0.08752   0.08086  -0.0001   1.0000   0.3397
  -5.000  -0.3449   0.08576   0.07920   0.0032   1.0000   0.3553
  -4.750  -0.3395   0.08315   0.07669   0.0065   1.0000   0.3692
  -4.500  -0.3710   0.08274   0.07644   0.0094   1.0000   0.3841
  -4.250  -0.3912   0.08275   0.07655   0.0119   1.0000   0.4043
  -4.000  -0.3889   0.08061   0.07451   0.0165   1.0000   0.4266
  -3.750  -0.3887   0.07839   0.07241   0.0212   1.0000   0.4508
  -3.500  -0.3635   0.07522   0.06932   0.0270   1.0000   0.4807
  -3.250  -0.3763   0.07458   0.06879   0.0322   1.0000   0.5157
  -3.000   0.1403   0.04804   0.04154  -0.0231   1.0000   1.0000
  -2.750   0.1429   0.04703   0.04065  -0.0220   1.0000   1.0000
  -2.500   0.1435   0.04619   0.03994  -0.0206   1.0000   1.0000
  -2.250   0.1417   0.04555   0.03943  -0.0189   1.0000   1.0000
  -2.000   0.0747   0.04837   0.04247  -0.0032   1.0000   0.9767
  -1.750  -0.0018   0.05111   0.04543   0.0130   1.0000   0.9427
  -1.500  -0.0748   0.05304   0.04756   0.0269   1.0000   0.9058
  -1.250  -0.1504   0.05464   0.04933   0.0404   1.0000   0.8765
  -1.000  -0.1710   0.04511   0.03686  -0.0290   1.0000   0.2472
  -0.750  -0.1492   0.04298   0.03476  -0.0297   1.0000   0.2383
  -0.500  -0.1188   0.04189   0.03310  -0.0318   1.0000   0.2242
  -0.250  -0.0803   0.04058   0.03161  -0.0351   0.9952   0.2188
   0.000  -0.0204   0.03959   0.03006  -0.0416   0.9839   0.2090
   0.250   0.0327   0.03910   0.02924  -0.0468   0.9717   0.2078
   0.500   0.0826   0.03855   0.02858  -0.0514   0.9586   0.2096
   0.750   0.1318   0.03820   0.02808  -0.0556   0.9442   0.2104
   1.000   0.1830   0.03787   0.02773  -0.0598   0.9280   0.2155
   1.250   0.2394   0.03753   0.02729  -0.0641   0.9093   0.2228
   1.500   0.2862   0.03706   0.02679  -0.0666   0.8879   0.2296
   1.750   0.3343   0.03649   0.02633  -0.0690   0.8666   0.2438
   2.000   0.3825   0.03575   0.02578  -0.0712   0.8459   0.2620
   2.250   0.4311   0.03472   0.02516  -0.0732   0.8256   0.3033
   2.500   0.4832   0.03189   0.02426  -0.0739   0.8071   1.0000
   2.750   0.5326   0.03136   0.02331  -0.0747   0.7856   1.0000
   3.000   0.5772   0.03066   0.02242  -0.0747   0.7628   1.0000
   3.250   0.6212   0.02974   0.02139  -0.0743   0.7388   1.0000
   3.500   0.6656   0.02857   0.02011  -0.0736   0.7135   1.0000
   3.750   0.7060   0.02745   0.01888  -0.0721   0.6849   1.0000
   4.000   0.7385   0.02669   0.01799  -0.0699   0.6503   1.0000
   4.250   0.7682   0.02615   0.01727  -0.0674   0.6116   1.0000
   4.500   0.7964   0.02590   0.01677  -0.0650   0.5714   1.0000
   4.750   0.8220   0.02608   0.01666  -0.0629   0.5328   1.0000
   5.000   0.8472   0.02657   0.01687  -0.0611   0.4994   1.0000
   5.250   0.8716   0.02724   0.01730  -0.0596   0.4706   1.0000
   5.500   0.8978   0.02800   0.01780  -0.0584   0.4470   1.0000
   5.750   0.9217   0.02894   0.01859  -0.0572   0.4260   1.0000
   6.000   0.9465   0.02993   0.01945  -0.0562   0.4076   1.0000
   6.250   0.9683   0.03106   0.02057  -0.0549   0.3904   1.0000
   6.500   0.9894   0.03233   0.02188  -0.0537   0.3748   1.0000
   6.750   1.0109   0.03363   0.02318  -0.0524   0.3600   1.0000
   7.000   1.0335   0.03483   0.02431  -0.0513   0.3452   1.0000
   7.250   1.0553   0.03586   0.02523  -0.0499   0.3283   1.0000
   7.500   1.0758   0.03684   0.02608  -0.0484   0.3104   1.0000
   7.750   1.0959   0.03790   0.02701  -0.0470   0.2927   1.0000
   8.000   1.1172   0.03913   0.02810  -0.0457   0.2757   1.0000
   8.250   1.1324   0.04063   0.02969  -0.0438   0.2594   1.0000
   8.500   1.1460   0.04242   0.03159  -0.0417   0.2434   1.0000
   8.750   1.1583   0.04450   0.03377  -0.0396   0.2282   1.0000
   9.000   1.1693   0.04683   0.03622  -0.0375   0.2146   1.0000
   9.250   1.1857   0.04898   0.03830  -0.0360   0.2026   1.0000
   9.500   1.1942   0.05141   0.04091  -0.0338   0.1928   1.0000
   9.750   1.1978   0.05450   0.04421  -0.0315   0.1858   1.0000
  10.000   1.1914   0.05799   0.04806  -0.0285   0.1804   1.0000
  10.250   1.2206   0.05995   0.04976  -0.0285   0.1737   1.0000
  10.500   1.1953   0.06462   0.05496  -0.0246   0.1721   1.0000
  10.750   1.1646   0.06966   0.06039  -0.0213   0.1713   1.0000
  11.000   1.1254   0.07514   0.06613  -0.0184   0.1715   1.0000
  11.250   1.0798   0.08234   0.07355  -0.0180   0.1726   1.0000
  11.500   1.0358   0.09133   0.08269  -0.0203   0.1738   1.0000
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