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=1,000,000 Ncrit=9


Details Polar file
Airfoil: PT40 (pt40-il)
Reynolds number: 1,000,000
Max Cl/Cd: 95 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-pt40-il-1000000.txt
Download as CSV file: xf-pt40-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3687   0.08808   0.08617  -0.0319   1.0000   0.0402
  -8.250  -0.3671   0.08582   0.08393  -0.0317   1.0000   0.0405
  -8.000  -0.5591   0.02015   0.01584  -0.0786   0.9841   0.0431
  -7.750  -0.5281   0.01804   0.01345  -0.0804   0.9828   0.0436
  -7.500  -0.4985   0.01728   0.01265  -0.0809   0.9790   0.0439
  -7.250  -0.4655   0.01679   0.01214  -0.0820   0.9761   0.0442
  -7.000  -0.4321   0.01633   0.01167  -0.0831   0.9734   0.0445
  -6.750  -0.3986   0.01587   0.01118  -0.0842   0.9709   0.0448
  -6.500  -0.3708   0.01553   0.01082  -0.0840   0.9640   0.0452
  -6.250  -0.3402   0.01507   0.01031  -0.0844   0.9585   0.0456
  -6.000  -0.3128   0.01444   0.00959  -0.0842   0.9500   0.0461
  -5.750  -0.2834   0.01378   0.00882  -0.0843   0.9420   0.0466
  -5.500  -0.2563   0.01324   0.00817  -0.0839   0.9299   0.0471
  -5.250  -0.2286   0.01285   0.00767  -0.0836   0.9171   0.0474
  -5.000  -0.2019   0.01215   0.00685  -0.0831   0.9034   0.0479
  -4.750  -0.1750   0.01170   0.00635  -0.0827   0.8888   0.0483
  -4.250  -0.1203   0.01127   0.00583  -0.0818   0.8598   0.0491
  -4.000  -0.0928   0.01108   0.00558  -0.0814   0.8466   0.0496
  -3.750  -0.0654   0.01092   0.00535  -0.0810   0.8331   0.0502
  -3.500  -0.0377   0.01072   0.00511  -0.0806   0.8208   0.0508
  -3.250  -0.0100   0.01051   0.00482  -0.0802   0.8093   0.0514
  -3.000   0.0177   0.01033   0.00457  -0.0799   0.7973   0.0519
  -2.750   0.0452   0.01007   0.00424  -0.0794   0.7840   0.0524
  -2.500   0.0718   0.00974   0.00387  -0.0789   0.7677   0.0531
  -2.000   0.1268   0.00951   0.00356  -0.0781   0.7349   0.0543
  -1.750   0.1544   0.00940   0.00341  -0.0777   0.7192   0.0551
  -1.500   0.1821   0.00930   0.00326  -0.0773   0.7045   0.0558
  -1.250   0.2100   0.00924   0.00314  -0.0769   0.6908   0.0567
  -1.000   0.2380   0.00913   0.00299  -0.0766   0.6774   0.0574
  -0.750   0.2652   0.00891   0.00275  -0.0762   0.6639   0.0584
  -0.500   0.2930   0.00885   0.00266  -0.0758   0.6486   0.0593
  -0.250   0.3209   0.00879   0.00258  -0.0755   0.6326   0.0603
   0.000   0.3487   0.00875   0.00250  -0.0752   0.6168   0.0613
   0.250   0.3764   0.00875   0.00244  -0.0748   0.5988   0.0623
   0.500   0.4037   0.00867   0.00231  -0.0744   0.5797   0.0635
   0.750   0.4311   0.00867   0.00227  -0.0740   0.5611   0.0650
   1.000   0.4586   0.00870   0.00225  -0.0736   0.5411   0.0664
   1.250   0.4860   0.00876   0.00223  -0.0732   0.5198   0.0677
   1.500   0.5133   0.00881   0.00222  -0.0728   0.4980   0.0689
   1.750   0.5404   0.00883   0.00218  -0.0723   0.4752   0.0710
   2.000   0.5672   0.00896   0.00222  -0.0719   0.4481   0.0729
   2.250   0.5935   0.00915   0.00228  -0.0713   0.4117   0.0749
   2.500   0.6185   0.00942   0.00234  -0.0706   0.3599   0.0775
   2.750   0.6427   0.00982   0.00249  -0.0698   0.3033   0.0802
   3.000   0.6681   0.01011   0.00263  -0.0691   0.2676   0.0830
   3.250   0.6940   0.01033   0.00275  -0.0685   0.2418   0.0884
   3.500   0.7204   0.01050   0.00286  -0.0681   0.2233   0.0959
   3.750   0.7472   0.01059   0.00297  -0.0676   0.2105   0.1177
   4.000   0.7744   0.00878   0.00330  -0.0679   0.1995   0.9987
   4.250   0.8009   0.00901   0.00343  -0.0674   0.1886   1.0000
   4.500   0.8272   0.00919   0.00355  -0.0668   0.1807   1.0000
   4.750   0.8533   0.00940   0.00370  -0.0662   0.1724   1.0000
   5.000   0.8795   0.00960   0.00383  -0.0657   0.1649   1.0000
   5.250   0.9057   0.00981   0.00399  -0.0651   0.1573   1.0000
   5.500   0.9318   0.01003   0.00415  -0.0646   0.1502   1.0000
   5.750   0.9580   0.01024   0.00431  -0.0640   0.1443   1.0000
   6.000   0.9840   0.01045   0.00448  -0.0635   0.1381   1.0000
   6.250   1.0100   0.01068   0.00467  -0.0629   0.1320   1.0000
   6.500   1.0356   0.01093   0.00486  -0.0623   0.1249   1.0000
   6.750   1.0612   0.01117   0.00506  -0.0617   0.1176   1.0000
   7.000   1.0862   0.01146   0.00529  -0.0611   0.1083   1.0000
   7.250   1.1100   0.01186   0.00557  -0.0602   0.0942   1.0000
   7.500   1.1311   0.01248   0.00601  -0.0590   0.0688   1.0000
   7.750   1.1519   0.01313   0.00652  -0.0577   0.0529   1.0000
   8.000   1.1747   0.01359   0.00693  -0.0567   0.0479   1.0000
   8.250   1.1976   0.01403   0.00733  -0.0557   0.0450   1.0000
   8.500   1.2210   0.01441   0.00771  -0.0548   0.0433   1.0000
   8.750   1.2442   0.01479   0.00808  -0.0539   0.0419   1.0000
   9.000   1.2664   0.01523   0.00852  -0.0528   0.0407   1.0000
   9.250   1.2880   0.01571   0.00900  -0.0517   0.0396   1.0000
   9.500   1.3107   0.01607   0.00938  -0.0507   0.0390   1.0000
   9.750   1.3327   0.01647   0.00980  -0.0496   0.0383   1.0000
  10.000   1.3540   0.01690   0.01024  -0.0485   0.0376   1.0000
  10.250   1.3744   0.01737   0.01072  -0.0472   0.0370   1.0000
  10.500   1.3934   0.01792   0.01127  -0.0457   0.0364   1.0000
  10.750   1.4091   0.01857   0.01195  -0.0438   0.0358   1.0000
  11.000   1.4252   0.01912   0.01253  -0.0418   0.0354   1.0000
  11.250   1.4422   0.01961   0.01306  -0.0400   0.0351   1.0000
  11.500   1.4585   0.02015   0.01364  -0.0382   0.0348   1.0000
  11.750   1.4742   0.02074   0.01426  -0.0364   0.0345   1.0000
  12.000   1.4893   0.02136   0.01492  -0.0346   0.0341   1.0000
  12.250   1.5038   0.02204   0.01563  -0.0328   0.0337   1.0000
  12.500   1.5174   0.02279   0.01641  -0.0309   0.0334   1.0000
  12.750   1.5301   0.02362   0.01728  -0.0291   0.0331   1.0000
  13.000   1.5415   0.02455   0.01825  -0.0273   0.0328   1.0000
  13.250   1.5515   0.02563   0.01937  -0.0255   0.0326   1.0000
  13.500   1.5592   0.02690   0.02070  -0.0236   0.0323   1.0000
  13.750   1.5644   0.02844   0.02229  -0.0218   0.0320   1.0000
  14.000   1.5657   0.03037   0.02430  -0.0199   0.0317   1.0000
  14.250   1.5741   0.03181   0.02580  -0.0187   0.0316   1.0000
  14.500   1.5826   0.03331   0.02736  -0.0176   0.0315   1.0000
  14.750   1.5899   0.03496   0.02908  -0.0167   0.0314   1.0000
  15.000   1.5963   0.03675   0.03094  -0.0158   0.0312   1.0000
  15.250   1.6015   0.03871   0.03298  -0.0151   0.0311   1.0000
  15.500   1.6056   0.04085   0.03520  -0.0145   0.0309   1.0000
  15.750   1.6085   0.04317   0.03759  -0.0140   0.0308   1.0000
  16.000   1.6102   0.04568   0.04018  -0.0137   0.0306   1.0000
  16.250   1.6110   0.04838   0.04296  -0.0135   0.0305   1.0000
  16.500   1.6105   0.05128   0.04594  -0.0135   0.0304   1.0000
  16.750   1.6091   0.05439   0.04914  -0.0137   0.0302   1.0000
  17.000   1.6069   0.05771   0.05254  -0.0141   0.0301   1.0000
  17.250   1.6039   0.06121   0.05613  -0.0147   0.0300   1.0000
  17.500   1.6001   0.06489   0.05989  -0.0154   0.0298   1.0000
  17.750   1.5955   0.06878   0.06386  -0.0164   0.0297   1.0000
  18.000   1.5900   0.07287   0.06804  -0.0174   0.0296   1.0000
  18.250   1.5837   0.07714   0.07239  -0.0187   0.0294   1.0000
  18.500   1.5763   0.08159   0.07693  -0.0201   0.0293   1.0000
  18.750   1.5684   0.08619   0.08162  -0.0216   0.0292   1.0000
  19.000   1.5602   0.09090   0.08640  -0.0232   0.0291   1.0000
  19.250   1.5518   0.09565   0.09124  -0.0249   0.0290   1.0000
<< Back to PT40 (pt40-il)

Polar data table (+)

Polar graphs


<< Back to PT40 (pt40-il)