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PT40 (pt40-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: PT40 (pt40-il)
Reynolds number: 50,000
Max Cl/Cd: 33.41 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-pt40-il-50000-n5.txt
Download as CSV file: xf-pt40-il-50000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3193   0.10420   0.09622  -0.0238   1.0000   0.1604
  -7.500  -0.3402   0.10327   0.09544  -0.0252   1.0000   0.1665
  -7.250  -0.3798   0.10288   0.09524  -0.0277   1.0000   0.1679
  -7.000  -0.3316   0.09721   0.08957  -0.0231   1.0000   0.1720
  -6.750  -0.3264   0.09484   0.08728  -0.0216   1.0000   0.1760
  -6.500  -0.3363   0.09311   0.08569  -0.0203   1.0000   0.1805
  -6.250  -0.3847   0.09256   0.08530  -0.0251   1.0000   0.1866
  -6.000  -0.3668   0.08878   0.08162  -0.0207   1.0000   0.1886
  -5.000  -0.3854   0.06877   0.06102  -0.0317   1.0000   0.1107
  -4.750  -0.3787   0.06650   0.05881  -0.0301   1.0000   0.1100
  -4.500  -0.3718   0.06407   0.05638  -0.0292   1.0000   0.1090
  -4.250  -0.3423   0.06031   0.05249  -0.0333   0.9936   0.1069
  -4.000  -0.3095   0.05584   0.04772  -0.0386   0.9855   0.1036
  -3.750  -0.2748   0.05134   0.04270  -0.0439   0.9773   0.1014
  -3.500  -0.2397   0.04875   0.03993  -0.0473   0.9702   0.1025
  -3.250  -0.2080   0.04627   0.03723  -0.0499   0.9612   0.1032
  -3.000  -0.1732   0.04371   0.03433  -0.0528   0.9531   0.1031
  -2.750  -0.1377   0.04138   0.03165  -0.0554   0.9450   0.1028
  -2.500  -0.1047   0.03959   0.02963  -0.0572   0.9362   0.1037
  -2.250  -0.0674   0.03798   0.02779  -0.0596   0.9285   0.1057
  -2.000  -0.0352   0.03657   0.02610  -0.0609   0.9190   0.1073
  -1.750   0.0039   0.03509   0.02433  -0.0631   0.9116   0.1082
  -1.500   0.0361   0.03396   0.02293  -0.0640   0.9017   0.1092
  -1.250   0.0753   0.03283   0.02161  -0.0660   0.8943   0.1117
  -1.000   0.1071   0.03198   0.02079  -0.0669   0.8843   0.1142
  -0.750   0.1456   0.03107   0.01980  -0.0686   0.8765   0.1165
  -0.500   0.1794   0.03034   0.01897  -0.0693   0.8669   0.1186
  -0.250   0.2154   0.02970   0.01818  -0.0702   0.8577   0.1221
   0.000   0.2484   0.02902   0.01760  -0.0708   0.8480   0.1256
   0.250   0.2826   0.02843   0.01704  -0.0713   0.8379   0.1291
   0.500   0.3159   0.02798   0.01655  -0.0715   0.8273   0.1334
   0.750   0.3512   0.02738   0.01600  -0.0721   0.8163   0.1389
   1.000   0.3819   0.02683   0.01550  -0.0716   0.8005   0.1447
   1.250   0.4151   0.02617   0.01481  -0.0711   0.7834   0.1517
   1.500   0.4472   0.02542   0.01415  -0.0706   0.7653   0.1620
   1.750   0.4787   0.02477   0.01357  -0.0699   0.7476   0.1748
   2.000   0.5084   0.02418   0.01314  -0.0692   0.7302   0.1998
   2.500   0.5678   0.02162   0.01275  -0.0672   0.6914   1.0000
   2.750   0.5937   0.02166   0.01257  -0.0658   0.6707   1.0000
   3.000   0.6194   0.02170   0.01243  -0.0644   0.6488   1.0000
   3.250   0.6446   0.02177   0.01232  -0.0630   0.6254   1.0000
   3.500   0.6674   0.02193   0.01237  -0.0613   0.5988   1.0000
   3.750   0.6914   0.02207   0.01234  -0.0597   0.5708   1.0000
   4.000   0.7148   0.02226   0.01235  -0.0581   0.5408   1.0000
   4.250   0.7374   0.02253   0.01243  -0.0564   0.5084   1.0000
   4.500   0.7588   0.02291   0.01261  -0.0547   0.4736   1.0000
   4.750   0.7799   0.02338   0.01285  -0.0530   0.4394   1.0000
   5.000   0.8005   0.02396   0.01315  -0.0513   0.4073   1.0000
   5.250   0.8210   0.02462   0.01362  -0.0498   0.3783   1.0000
   5.500   0.8418   0.02534   0.01412  -0.0484   0.3547   1.0000
   5.750   0.8629   0.02611   0.01470  -0.0471   0.3348   1.0000
   6.000   0.8849   0.02690   0.01535  -0.0460   0.3177   1.0000
   6.250   0.9073   0.02772   0.01606  -0.0450   0.3028   1.0000
   6.500   0.9301   0.02857   0.01679  -0.0441   0.2900   1.0000
   6.750   0.9531   0.02943   0.01756  -0.0432   0.2780   1.0000
   7.000   0.9758   0.03033   0.01846  -0.0423   0.2665   1.0000
   7.250   0.9991   0.03124   0.01925  -0.0415   0.2563   1.0000
   7.500   1.0196   0.03217   0.02023  -0.0404   0.2445   1.0000
   7.750   1.0381   0.03309   0.02114  -0.0390   0.2321   1.0000
   8.000   1.0552   0.03397   0.02192  -0.0375   0.2197   1.0000
   8.250   1.0702   0.03485   0.02280  -0.0358   0.2072   1.0000
   8.500   1.0844   0.03583   0.02387  -0.0341   0.1954   1.0000
   8.750   1.0997   0.03683   0.02485  -0.0325   0.1854   1.0000
   9.000   1.1138   0.03785   0.02588  -0.0308   0.1757   1.0000
   9.250   1.1263   0.03904   0.02720  -0.0290   0.1657   1.0000
   9.500   1.1384   0.04020   0.02835  -0.0272   0.1571   1.0000
   9.750   1.1484   0.04144   0.02965  -0.0252   0.1485   1.0000
  10.000   1.1564   0.04288   0.03120  -0.0231   0.1402   1.0000
  10.250   1.1648   0.04418   0.03244  -0.0211   0.1336   1.0000
  10.500   1.1714   0.04593   0.03436  -0.0191   0.1264   1.0000
  10.750   1.1786   0.04759   0.03604  -0.0174   0.1208   1.0000
  11.000   1.1839   0.04948   0.03805  -0.0157   0.1154   1.0000
  11.250   1.1902   0.05132   0.03991  -0.0142   0.1113   1.0000
  11.500   1.1931   0.05374   0.04254  -0.0128   0.1071   1.0000
  11.750   1.1986   0.05571   0.04453  -0.0115   0.1041   1.0000
  12.000   1.1993   0.05842   0.04740  -0.0105   0.1013   1.0000
  12.250   1.1954   0.06166   0.05088  -0.0097   0.0986   1.0000
  12.500   1.1940   0.06454   0.05387  -0.0091   0.0964   1.0000
  12.750   1.1983   0.06679   0.05610  -0.0085   0.0945   1.0000
  13.000   1.1817   0.07186   0.06149  -0.0090   0.0930   1.0000
  13.250   1.1586   0.07807   0.06804  -0.0105   0.0918   1.0000
  13.500   1.1266   0.08613   0.07640  -0.0137   0.0910   1.0000
  13.750   1.0621   0.10108   0.09179  -0.0220   0.0910   1.0000
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