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USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 5 AIRFOIL (usa5-il)
Reynolds number: 200,000
Max Cl/Cd: 72.37 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa5-il-200000.txt
Download as CSV file: xf-usa5-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 5 AIRFOIL                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.250  -0.3575   0.08076   0.07779  -0.0179   1.0000   0.0327
  -6.000  -0.3582   0.07847   0.07553  -0.0173   1.0000   0.0338
  -5.750  -0.3576   0.07612   0.07320  -0.0171   1.0000   0.0347
  -5.500  -0.3553   0.07367   0.07077  -0.0172   1.0000   0.0357
  -5.250  -0.3507   0.07104   0.06813  -0.0177   1.0000   0.0371
  -5.000  -0.3414   0.06840   0.06547  -0.0194   1.0000   0.0389
  -4.750  -0.2736   0.06393   0.06067  -0.0347   0.9940   0.0408
  -4.500  -0.2501   0.05686   0.05358  -0.0390   0.9879   0.0420
  -4.250  -0.2172   0.05294   0.04964  -0.0427   0.9832   0.0440
  -4.000  -0.1823   0.04932   0.04590  -0.0470   0.9761   0.0476
  -3.750  -0.1248   0.04419   0.04026  -0.0556   0.9705   0.0538
  -3.500  -0.1002   0.04067   0.03680  -0.0572   0.9635   0.0564
  -2.500   0.0493   0.02103   0.01498  -0.0644   0.9417   0.0403
  -2.250   0.0764   0.01811   0.01153  -0.0639   0.9335   0.0431
  -2.000   0.1171   0.01678   0.00994  -0.0660   0.9296   0.0497
  -1.750   0.1547   0.01515   0.00799  -0.0673   0.9243   0.0566
  -1.500   0.1908   0.01469   0.00748  -0.0685   0.9168   0.0703
  -1.250   0.2255   0.01433   0.00705  -0.0695   0.9086   0.0865
  -1.000   0.2618   0.01371   0.00629  -0.0706   0.9010   0.1052
  -0.750   0.2912   0.01307   0.00570  -0.0706   0.8899   0.1257
  -0.500   0.3217   0.01255   0.00516  -0.0708   0.8784   0.1411
  -0.250   0.3519   0.01212   0.00472  -0.0708   0.8648   0.1552
   0.000   0.3812   0.01175   0.00435  -0.0706   0.8485   0.1698
   0.250   0.4105   0.01135   0.00390  -0.0704   0.8307   0.1800
   0.500   0.4363   0.01108   0.00361  -0.0695   0.8082   0.1912
   0.750   0.4626   0.01084   0.00333  -0.0688   0.7859   0.2034
   1.000   0.4875   0.01062   0.00311  -0.0678   0.7622   0.2233
   1.250   0.5988   0.00885   0.00285  -0.0864   0.7244   1.0000
   1.500   0.6221   0.00903   0.00279  -0.0850   0.6900   1.0000
   1.750   0.6445   0.00927   0.00278  -0.0836   0.6530   1.0000
   2.000   0.6665   0.00956   0.00284  -0.0822   0.6155   1.0000
   2.250   0.6884   0.00988   0.00292  -0.0808   0.5797   1.0000
   2.500   0.7106   0.01018   0.00303  -0.0795   0.5460   1.0000
   2.750   0.7332   0.01049   0.00316  -0.0784   0.5195   1.0000
   3.000   0.7561   0.01080   0.00333  -0.0774   0.4985   1.0000
   3.250   0.7794   0.01112   0.00355  -0.0765   0.4814   1.0000
   3.500   0.8030   0.01143   0.00376  -0.0756   0.4673   1.0000
   3.750   0.8268   0.01174   0.00400  -0.0748   0.4552   1.0000
   4.000   0.8510   0.01202   0.00427  -0.0741   0.4441   1.0000
   4.250   0.8751   0.01232   0.00456  -0.0734   0.4338   1.0000
   4.500   0.8986   0.01262   0.00487  -0.0725   0.4212   1.0000
   4.750   0.9192   0.01289   0.00505  -0.0711   0.3953   1.0000
   5.000   0.9416   0.01314   0.00530  -0.0701   0.3775   1.0000
   5.250   0.9640   0.01339   0.00555  -0.0691   0.3609   1.0000
   5.500   0.9856   0.01364   0.00579  -0.0679   0.3377   1.0000
   5.750   1.0067   0.01391   0.00606  -0.0666   0.3065   1.0000
   6.000   1.0270   0.01430   0.00632  -0.0652   0.2570   1.0000
   6.250   1.0414   0.01536   0.00685  -0.0631   0.1738   1.0000
   6.500   1.0531   0.01687   0.00778  -0.0607   0.0796   1.0000
   6.750   1.0642   0.01843   0.00901  -0.0579   0.0295   1.0000
   7.000   1.0827   0.01922   0.00996  -0.0561   0.0267   1.0000
   7.250   1.1006   0.02006   0.01102  -0.0542   0.0253   1.0000
   7.500   1.1166   0.02103   0.01218  -0.0521   0.0244   1.0000
   7.750   1.1300   0.02214   0.01348  -0.0496   0.0239   1.0000
   8.000   1.1361   0.02377   0.01525  -0.0461   0.0225   1.0000
   8.250   1.1418   0.02543   0.01699  -0.0426   0.0215   1.0000
   8.500   1.1508   0.02694   0.01859  -0.0397   0.0214   1.0000
   8.750   1.1612   0.02867   0.02038  -0.0370   0.0214   1.0000
   9.000   1.1763   0.03082   0.02258  -0.0350   0.0215   1.0000
   9.250   1.1987   0.03304   0.02486  -0.0341   0.0221   1.0000
   9.500   1.2181   0.03418   0.02618  -0.0325   0.0229   1.0000
   9.750   1.2423   0.03597   0.02835  -0.0309   0.0258   1.0000
  10.000   1.2828   0.04198   0.03457  -0.0324   0.0315   1.0000
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