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

PT40 (pt40-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: PT40 (pt40-il)
Reynolds number: 100,000
Max Cl/Cd: 48.75 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pt40-il-100000.txt
Download as CSV file: xf-pt40-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3209   0.09900   0.09334  -0.0267   1.0000   0.1197
  -7.500  -0.3629   0.09834   0.09286  -0.0296   1.0000   0.1222
  -7.250  -0.3786   0.09524   0.08987  -0.0304   1.0000   0.1231
  -7.000  -0.3386   0.09152   0.08613  -0.0258   1.0000   0.1253
  -6.750  -0.3347   0.08939   0.08408  -0.0239   1.0000   0.1275
  -6.500  -0.3440   0.08767   0.08247  -0.0218   1.0000   0.1297
  -6.250  -0.3595   0.08584   0.08074  -0.0211   1.0000   0.1324
  -6.000  -0.4141   0.08384   0.07864  -0.0290   1.0000   0.1363
  -5.750  -0.4045   0.08066   0.07563  -0.0243   1.0000   0.1372
  -5.500  -0.4001   0.07860   0.07367  -0.0204   1.0000   0.1384
  -5.250  -0.3992   0.07684   0.07198  -0.0174   1.0000   0.1401
  -5.000  -0.4003   0.07502   0.07021  -0.0156   1.0000   0.1426
  -4.750  -0.4100   0.07182   0.06669  -0.0235   1.0000   0.1519
  -4.500  -0.4043   0.06909   0.06411  -0.0203   1.0000   0.1531
  -4.250  -0.3949   0.06706   0.06218  -0.0182   0.9992   0.1552
  -4.000  -0.3535   0.06308   0.05793  -0.0282   0.9916   0.1697
  -3.750  -0.3103   0.06209   0.05667  -0.0349   0.9836   0.1859
  -3.500  -0.2863   0.05702   0.05182  -0.0359   0.9775   0.1894
  -3.250  -0.2518   0.05467   0.04956  -0.0381   0.9713   0.1971
  -3.000  -0.2102   0.05370   0.04813  -0.0455   0.9610   0.2270
  -2.750  -0.1802   0.04936   0.04411  -0.0464   0.9564   0.2324
  -2.500  -0.1517   0.04742   0.04217  -0.0482   0.9467   0.2536
  -2.250  -0.1180   0.04551   0.04032  -0.0501   0.9396   0.2781
  -2.000  -0.0878   0.04400   0.03881  -0.0517   0.9304   0.3176
  -1.250   0.0869   0.03225   0.02431  -0.0713   0.9108   0.1517
  -1.000   0.1173   0.03036   0.02242  -0.0720   0.9001   0.1470
  -0.750   0.1660   0.02859   0.02042  -0.0753   0.8945   0.1430
  -0.500   0.2031   0.02754   0.01919  -0.0765   0.8850   0.1426
  -0.250   0.2499   0.02622   0.01769  -0.0791   0.8776   0.1414
   0.000   0.2940   0.02490   0.01626  -0.0806   0.8664   0.1420
   0.250   0.3440   0.02332   0.01463  -0.0826   0.8552   0.1460
   0.500   0.3770   0.02241   0.01367  -0.0820   0.8396   0.1484
   0.750   0.4119   0.02163   0.01285  -0.0819   0.8273   0.1508
   1.000   0.4461   0.02054   0.01192  -0.0819   0.8152   0.1564
   1.250   0.4732   0.02002   0.01148  -0.0807   0.7996   0.1621
   1.500   0.5027   0.01950   0.01097  -0.0797   0.7852   0.1676
   1.750   0.5334   0.01874   0.01032  -0.0789   0.7716   0.1781
   2.000   0.5588   0.01837   0.01000  -0.0774   0.7538   0.1910
   2.250   0.5851   0.01794   0.00970  -0.0761   0.7359   0.2169
   2.500   0.6252   0.01561   0.00940  -0.0765   0.7169   1.0000
   2.750   0.6510   0.01560   0.00914  -0.0749   0.6966   1.0000
   3.000   0.6762   0.01560   0.00895  -0.0733   0.6747   1.0000
   3.250   0.7010   0.01559   0.00877  -0.0717   0.6505   1.0000
   3.500   0.7251   0.01562   0.00863  -0.0700   0.6232   1.0000
   3.750   0.7477   0.01570   0.00855  -0.0681   0.5900   1.0000
   4.000   0.7695   0.01586   0.00849  -0.0660   0.5503   1.0000
   4.250   0.7898   0.01620   0.00850  -0.0638   0.5033   1.0000
   4.500   0.8089   0.01674   0.00869  -0.0616   0.4521   1.0000
   4.750   0.8285   0.01742   0.00899  -0.0597   0.4094   1.0000
   5.000   0.8494   0.01814   0.00942  -0.0581   0.3763   1.0000
   5.250   0.8715   0.01890   0.00989  -0.0569   0.3526   1.0000
   5.500   0.8947   0.01966   0.01044  -0.0558   0.3334   1.0000
   5.750   0.9184   0.02040   0.01105  -0.0549   0.3171   1.0000
   6.000   0.9420   0.02115   0.01168  -0.0540   0.3021   1.0000
   6.250   0.9651   0.02190   0.01231  -0.0530   0.2878   1.0000
   6.500   0.9876   0.02263   0.01286  -0.0520   0.2733   1.0000
   6.750   1.0093   0.02326   0.01339  -0.0509   0.2593   1.0000
   7.000   1.0308   0.02391   0.01407  -0.0498   0.2469   1.0000
   7.250   1.0537   0.02471   0.01477  -0.0489   0.2363   1.0000
   7.500   1.0754   0.02536   0.01540  -0.0478   0.2254   1.0000
   7.750   1.0962   0.02612   0.01620  -0.0466   0.2146   1.0000
   8.000   1.1177   0.02692   0.01687  -0.0456   0.2039   1.0000
   8.250   1.1360   0.02755   0.01754  -0.0440   0.1920   1.0000
   8.500   1.1531   0.02833   0.01837  -0.0423   0.1792   1.0000
   8.750   1.1691   0.02918   0.01922  -0.0405   0.1656   1.0000
   9.000   1.1838   0.03013   0.02012  -0.0385   0.1519   1.0000
   9.250   1.1985   0.03117   0.02113  -0.0366   0.1397   1.0000
   9.500   1.2145   0.03233   0.02224  -0.0349   0.1301   1.0000
   9.750   1.2340   0.03347   0.02317  -0.0338   0.1231   1.0000
  10.000   1.2489   0.03482   0.02473  -0.0319   0.1167   1.0000
  10.250   1.2701   0.03606   0.02579  -0.0311   0.1121   1.0000
  10.500   1.2838   0.03754   0.02752  -0.0292   0.1077   1.0000
  10.750   1.3030   0.03871   0.02861  -0.0282   0.1043   1.0000
  11.000   1.3195   0.04044   0.03045  -0.0269   0.1014   1.0000
  11.250   1.3326   0.04226   0.03248  -0.0251   0.0989   1.0000
  11.500   1.3479   0.04392   0.03424  -0.0238   0.0968   1.0000
  11.750   1.3679   0.04553   0.03580  -0.0231   0.0950   1.0000
  12.000   1.3784   0.04779   0.03824  -0.0214   0.0936   1.0000
  12.250   1.3753   0.05023   0.04103  -0.0182   0.0926   1.0000
  12.500   1.3701   0.05277   0.04388  -0.0150   0.0915   1.0000
  12.750   1.3634   0.05544   0.04681  -0.0122   0.0906   1.0000
  13.000   1.3553   0.05829   0.04990  -0.0098   0.0898   1.0000
  13.250   1.3458   0.06135   0.05318  -0.0077   0.0891   1.0000
  13.500   1.3327   0.06486   0.05691  -0.0061   0.0886   1.0000
  13.750   1.3074   0.06962   0.06196  -0.0052   0.0885   1.0000
  14.000   1.2432   0.07896   0.07184  -0.0068   0.0895   1.0000
  14.250   1.1267   0.09899   0.09248  -0.0175   0.0924   1.0000
  14.500   1.0505   0.11846   0.11219  -0.0292   0.0944   1.0000
<< Back to PT40 (pt40-il)

Polar data table (+)

Polar graphs


<< Back to PT40 (pt40-il)