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EPPLER 657 AIRFOIL (e657-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 657 AIRFOIL (e657-il)
Reynolds number: 50,000
Max Cl/Cd: 7.55 at α=11°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e657-il-50000.txt
Download as CSV file: xf-e657-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 657 AIRFOIL                              
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.4768   0.12608   0.12098  -0.0112   1.0000   0.2543
  -7.000  -0.4605   0.12219   0.11710  -0.0091   1.0000   0.2643
  -6.750  -0.4992   0.12170   0.11674  -0.0075   1.0000   0.2716
  -6.500  -0.4796   0.11785   0.11287  -0.0053   1.0000   0.2845
  -6.250  -0.4896   0.11538   0.11046  -0.0031   1.0000   0.2953
  -6.000  -0.5153   0.11419   0.10937  -0.0004   1.0000   0.3048
  -5.750  -0.5086   0.11114   0.10634   0.0020   1.0000   0.3186
  -5.500  -0.5182   0.10873   0.10399   0.0048   1.0000   0.3311
  -4.250  -0.5414   0.06521   0.05835  -0.0371   1.0000   0.1229
  -4.000  -0.5241   0.06136   0.05437  -0.0374   1.0000   0.1196
  -3.750  -0.5000   0.05665   0.04906  -0.0392   1.0000   0.1138
  -3.500  -0.4706   0.05285   0.04421  -0.0407   1.0000   0.1081
  -3.250  -0.4459   0.05011   0.04095  -0.0411   1.0000   0.1081
  -3.000  -0.4233   0.04783   0.03848  -0.0412   1.0000   0.1120
  -2.750  -0.4002   0.04630   0.03670  -0.0411   1.0000   0.1177
  -2.500  -0.3740   0.04484   0.03464  -0.0409   1.0000   0.1240
  -2.250  -0.3516   0.04361   0.03338  -0.0406   1.0000   0.1365
  -2.000  -0.3284   0.04263   0.03229  -0.0400   1.0000   0.1534
  -1.750  -0.3064   0.04177   0.03158  -0.0392   1.0000   0.1798
  -1.500  -0.2845   0.04103   0.03101  -0.0381   1.0000   0.2248
  -1.250  -0.2567   0.03960   0.03066  -0.0382   1.0000   0.3640
  -1.000  -0.2743   0.03844   0.03209  -0.0243   1.0000   0.8276
  -0.750  -0.2548   0.03790   0.03131  -0.0210   1.0000   1.0000
  -0.500  -0.2370   0.03824   0.03116  -0.0210   1.0000   1.0000
  -0.250  -0.2181   0.03873   0.03123  -0.0213   1.0000   1.0000
   0.000  -0.1986   0.03934   0.03147  -0.0216   1.0000   1.0000
   0.250  -0.1787   0.04004   0.03184  -0.0220   1.0000   1.0000
   0.500  -0.1587   0.04082   0.03233  -0.0225   1.0000   1.0000
   0.750  -0.1387   0.04166   0.03290  -0.0230   1.0000   1.0000
   1.000  -0.1188   0.04257   0.03355  -0.0235   1.0000   1.0000
   1.250  -0.0990   0.04353   0.03429  -0.0241   1.0000   1.0000
   1.500  -0.0794   0.04455   0.03510  -0.0246   1.0000   1.0000
   1.750  -0.0599   0.04562   0.03597  -0.0251   1.0000   1.0000
   2.000  -0.0407   0.04674   0.03692  -0.0257   1.0000   1.0000
   2.250  -0.0060   0.04919   0.03916  -0.0291   0.9939   1.0000
   2.500   0.0265   0.05135   0.04113  -0.0323   0.9848   1.0000
   3.000   0.0854   0.05532   0.04483  -0.0374   0.9639   1.0000
   3.250   0.1182   0.05793   0.04729  -0.0405   0.9546   1.0000
   3.500   0.1501   0.06022   0.04947  -0.0434   0.9420   1.0000
   3.750   0.1722   0.06154   0.05073  -0.0446   0.9286   1.0000
   4.000   0.1940   0.06310   0.05222  -0.0458   0.9157   1.0000
   4.250   0.2192   0.06521   0.05428  -0.0475   0.9049   1.0000
   4.500   0.2552   0.06829   0.05729  -0.0509   0.8933   1.0000
   4.750   0.2714   0.06927   0.05825  -0.0511   0.8790   1.0000
   5.000   0.2887   0.07076   0.05972  -0.0516   0.8660   1.0000
   5.250   0.3112   0.07292   0.06186  -0.0529   0.8550   1.0000
   5.500   0.3471   0.07616   0.06507  -0.0561   0.8433   1.0000
   5.750   0.3597   0.07715   0.06608  -0.0558   0.8288   1.0000
   6.000   0.3732   0.07869   0.06763  -0.0559   0.8160   1.0000
   6.250   0.3944   0.08101   0.06997  -0.0571   0.8051   1.0000
   6.500   0.4331   0.08461   0.07359  -0.0604   0.7926   1.0000
   6.750   0.4401   0.08541   0.07443  -0.0596   0.7782   1.0000
   7.000   0.4506   0.08704   0.07610  -0.0595   0.7653   1.0000
   7.250   0.4683   0.08937   0.07847  -0.0603   0.7542   1.0000
   7.500   0.4988   0.09240   0.08155  -0.0625   0.7417   1.0000
   7.750   0.5270   0.09499   0.08420  -0.0641   0.7262   1.0000
   8.000   0.5304   0.09608   0.08537  -0.0633   0.7123   1.0000
   8.250   0.5376   0.09802   0.08737  -0.0631   0.7002   1.0000
   8.500   0.5537   0.10031   0.08973  -0.0638   0.6871   1.0000
   8.750   0.5720   0.10260   0.09212  -0.0646   0.6729   1.0000
   9.000   0.6556   0.09791   0.08745  -0.0640   0.5930   1.0000
   9.250   0.6776   0.09925   0.08891  -0.0642   0.5770   1.0000
   9.500   0.6955   0.10094   0.09068  -0.0644   0.5627   1.0000
   9.750   0.7148   0.10250   0.09235  -0.0646   0.5478   1.0000
  10.000   0.7336   0.10422   0.09417  -0.0648   0.5340   1.0000
  10.250   0.7543   0.10573   0.09581  -0.0650   0.5195   1.0000
  10.500   0.7785   0.10707   0.09728  -0.0652   0.5057   1.0000
  10.750   0.8089   0.10791   0.09827  -0.0654   0.4917   1.0000
  11.000   0.8264   0.10947   0.09996  -0.0653   0.4778   1.0000
  11.250   0.8192   0.11324   0.10382  -0.0653   0.4651   1.0000
  11.500   0.8263   0.11593   0.10661  -0.0654   0.4520   1.0000
  11.750   0.8410   0.11807   0.10888  -0.0655   0.4392   1.0000
  12.000   0.8626   0.11938   0.11036  -0.0653   0.4256   1.0000
  12.250   0.8941   0.11948   0.11063  -0.0648   0.4120   1.0000
  12.500   0.8777   0.12502   0.11623  -0.0658   0.4006   1.0000
  12.750   0.8779   0.12879   0.12010  -0.0664   0.3890   1.0000
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