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EPPLER 654 AIRFOIL (e654-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 654 AIRFOIL (e654-il)
Reynolds number: 50,000
Max Cl/Cd: 5.12 at α=10°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e654-il-50000.txt
Download as CSV file: xf-e654-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 654 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
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.5004   0.12804   0.12330  -0.0051   1.0000   0.2979
  -6.500  -0.4672   0.12323   0.11846  -0.0036   1.0000   0.3059
  -6.250  -0.5263   0.12387   0.11927  -0.0007   1.0000   0.3160
  -6.000  -0.4831   0.11893   0.11424   0.0004   1.0000   0.3268
  -5.750  -0.4983   0.11676   0.11215   0.0028   1.0000   0.3381
  -5.500  -0.5475   0.11676   0.11228   0.0072   1.0000   0.3509
  -5.250  -0.5129   0.11252   0.10800   0.0080   1.0000   0.3641
  -5.000  -0.6087   0.07414   0.06809  -0.0402   1.0000   0.1297
  -4.750  -0.5942   0.07007   0.06396  -0.0402   1.0000   0.1239
  -4.500  -0.5690   0.06271   0.05553  -0.0450   1.0000   0.1125
  -4.250  -0.5477   0.05896   0.05138  -0.0463   1.0000   0.1108
  -4.000  -0.5236   0.05575   0.04761  -0.0477   1.0000   0.1110
  -3.750  -0.4968   0.05315   0.04426  -0.0491   1.0000   0.1137
  -3.500  -0.4744   0.05081   0.04177  -0.0495   1.0000   0.1170
  -3.250  -0.4516   0.04934   0.04010  -0.0496   1.0000   0.1234
  -3.000  -0.4264   0.04779   0.03814  -0.0499   1.0000   0.1314
  -2.750  -0.4030   0.04681   0.03695  -0.0498   1.0000   0.1429
  -2.500  -0.3803   0.04589   0.03598  -0.0495   1.0000   0.1589
  -2.250  -0.3586   0.04510   0.03532  -0.0488   1.0000   0.1809
  -2.000  -0.3371   0.04449   0.03487  -0.0480   1.0000   0.2148
  -1.500  -0.2979   0.04237   0.03592  -0.0436   1.0000   0.5916
  -1.250  -0.3082   0.04345   0.03721  -0.0329   1.0000   0.7467
  -1.000  -0.3177   0.04402   0.03772  -0.0233   1.0000   0.8131
  -0.750  -0.3271   0.04408   0.03770  -0.0142   1.0000   0.8690
  -0.500  -0.3282   0.04390   0.03741  -0.0071   1.0000   0.9321
  -0.250  -0.2513   0.04581   0.03885  -0.0175   1.0000   1.0000
   0.000  -0.2472   0.04525   0.03808  -0.0161   1.0000   1.0000
   0.250  -0.2370   0.04513   0.03771  -0.0157   1.0000   1.0000
   0.500  -0.2217   0.04541   0.03775  -0.0162   1.0000   1.0000
   0.750  -0.1899   0.04702   0.03905  -0.0198   0.9950   1.0000
   1.000  -0.1577   0.04865   0.04040  -0.0233   0.9879   1.0000
   1.250  -0.1223   0.05083   0.04229  -0.0274   0.9814   1.0000
   1.500  -0.0870   0.05283   0.04404  -0.0314   0.9722   1.0000
   1.750  -0.0592   0.05415   0.04517  -0.0339   0.9623   1.0000
   2.000  -0.0286   0.05603   0.04685  -0.0370   0.9539   1.0000
   2.250   0.0087   0.05852   0.04913  -0.0411   0.9443   1.0000
   2.500   0.0315   0.05956   0.05004  -0.0426   0.9326   1.0000
   2.750   0.0601   0.06156   0.05189  -0.0452   0.9239   1.0000
   3.000   0.0944   0.06388   0.05406  -0.0486   0.9129   1.0000
   3.250   0.1133   0.06490   0.05499  -0.0495   0.9011   1.0000
   3.500   0.1565   0.06883   0.05876  -0.0544   0.8941   1.0000
   3.750   0.1698   0.06900   0.05887  -0.0542   0.8803   1.0000
   4.000   0.1892   0.07053   0.06033  -0.0552   0.8698   1.0000
   4.250   0.2273   0.07375   0.06345  -0.0590   0.8604   1.0000
   4.500   0.2388   0.07440   0.06407  -0.0587   0.8477   1.0000
   4.750   0.2687   0.07734   0.06693  -0.0614   0.8401   1.0000
   5.000   0.2908   0.07891   0.06846  -0.0626   0.8269   1.0000
   5.250   0.3038   0.08027   0.06980  -0.0627   0.8158   1.0000
   5.500   0.3438   0.08409   0.07356  -0.0666   0.8072   1.0000
   5.750   0.3490   0.08453   0.07402  -0.0656   0.7940   1.0000
   6.000   0.3702   0.08705   0.07652  -0.0670   0.7858   1.0000
   6.250   0.3970   0.08945   0.07891  -0.0689   0.7735   1.0000
   6.500   0.4042   0.09076   0.08023  -0.0685   0.7623   1.0000
   6.750   0.4468   0.09518   0.08463  -0.0724   0.7542   1.0000
   7.000   0.4455   0.09542   0.08492  -0.0709   0.7409   1.0000
   7.250   0.4605   0.09782   0.08734  -0.0717   0.7322   1.0000
   7.500   0.4905   0.10080   0.09034  -0.0738   0.7208   1.0000
   7.750   0.4924   0.10211   0.09170  -0.0732   0.7095   1.0000
   8.000   0.5284   0.10628   0.09589  -0.0761   0.7017   1.0000
   8.250   0.5304   0.10719   0.09686  -0.0754   0.6887   1.0000
   8.500   0.5406   0.10964   0.09935  -0.0759   0.6802   1.0000
   8.750   0.5735   0.11311   0.10287  -0.0782   0.6690   1.0000
   9.000   0.5693   0.11440   0.10422  -0.0774   0.6584   1.0000
   9.250   0.5975   0.11819   0.10808  -0.0794   0.6504   1.0000
   9.500   0.6049   0.11985   0.10981  -0.0795   0.6380   1.0000
   9.750   0.6110   0.12237   0.11239  -0.0800   0.6296   1.0000
  10.000   0.6476   0.12647   0.11656  -0.0823   0.6187   1.0000
  10.250   0.6385   0.12756   0.11771  -0.0817   0.6084   1.0000
  10.500   0.6560   0.13092   0.12117  -0.0830   0.6007   1.0000
  10.750   0.6791   0.13387   0.12421  -0.0842   0.5882   1.0000
  11.000   0.6736   0.13585   0.12625  -0.0843   0.5798   1.0000
  11.250   0.6991   0.13962   0.13010  -0.0859   0.5707   1.0000
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