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EPPLER 655 AIRFOIL (e655-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 655 AIRFOIL (e655-il)
Reynolds number: 50,000
Max Cl/Cd: 4.91 at α=10°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e655-il-50000.txt
Download as CSV file: xf-e655-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 655 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  -0.4659   0.11774   0.11272  -0.0042   1.0000   0.3406
  -7.000  -0.6954   0.08635   0.08103  -0.0319   1.0000   0.1359
  -6.750  -0.6929   0.08210   0.07670  -0.0309   1.0000   0.1320
  -6.500  -0.7148   0.07265   0.06632  -0.0334   1.0000   0.1195
  -6.250  -0.7080   0.06829   0.06178  -0.0330   1.0000   0.1184
  -6.000  -0.6995   0.06393   0.05711  -0.0328   1.0000   0.1171
  -5.750  -0.6885   0.05950   0.05214  -0.0330   1.0000   0.1171
  -5.500  -0.6735   0.05564   0.04757  -0.0332   1.0000   0.1189
  -5.250  -0.6567   0.05249   0.04413  -0.0331   1.0000   0.1220
  -5.000  -0.6385   0.05030   0.04178  -0.0326   1.0000   0.1273
  -4.750  -0.6171   0.04792   0.03871  -0.0326   1.0000   0.1339
  -4.500  -0.5976   0.04597   0.03674  -0.0321   1.0000   0.1414
  -4.250  -0.5757   0.04427   0.03462  -0.0317   1.0000   0.1520
  -4.000  -0.5546   0.04301   0.03319  -0.0312   1.0000   0.1653
  -3.750  -0.5335   0.04183   0.03190  -0.0306   1.0000   0.1821
  -3.500  -0.5137   0.04076   0.03103  -0.0297   1.0000   0.2032
  -3.250  -0.4930   0.03989   0.03026  -0.0288   1.0000   0.2338
  -3.000  -0.4732   0.03904   0.02973  -0.0278   1.0000   0.2774
  -2.750  -0.4536   0.03795   0.02924  -0.0269   1.0000   0.3541
  -2.500  -0.4403   0.03675   0.02975  -0.0234   1.0000   0.5329
  -2.250  -0.4461   0.03782   0.03150  -0.0135   1.0000   0.6896
  -2.000  -0.4509   0.03897   0.03262  -0.0045   1.0000   0.7667
  -1.750  -0.4560   0.03973   0.03328   0.0043   1.0000   0.8226
  -1.500  -0.4598   0.04016   0.03360   0.0126   1.0000   0.8722
  -1.250  -0.4381   0.04122   0.03442   0.0162   1.0000   0.9311
  -1.000  -0.2184   0.04875   0.04083  -0.0201   1.0000   1.0000
  -0.750  -0.2191   0.04814   0.04007  -0.0179   1.0000   1.0000
  -0.500  -0.2191   0.04760   0.03939  -0.0157   1.0000   1.0000
  -0.250  -0.2185   0.04710   0.03875  -0.0135   1.0000   1.0000
   0.000  -0.2176   0.04664   0.03816  -0.0113   1.0000   1.0000
   0.250  -0.2160   0.04621   0.03761  -0.0091   1.0000   1.0000
   0.500  -0.2137   0.04584   0.03711  -0.0071   1.0000   1.0000
   0.750  -0.2104   0.04553   0.03668  -0.0052   1.0000   1.0000
   1.000  -0.2043   0.04541   0.03643  -0.0038   1.0000   1.0000
   1.250  -0.1942   0.04556   0.03646  -0.0031   1.0000   1.0000
   1.500  -0.1749   0.04640   0.03714  -0.0042   0.9975   1.0000
   1.750  -0.1429   0.04827   0.03884  -0.0077   0.9893   1.0000
   2.000  -0.1096   0.05047   0.04084  -0.0113   0.9813   1.0000
   2.250  -0.0750   0.05268   0.04290  -0.0152   0.9710   1.0000
   2.500  -0.0472   0.05410   0.04419  -0.0177   0.9591   1.0000
   2.750  -0.0210   0.05557   0.04555  -0.0199   0.9474   1.0000
   3.000   0.0067   0.05738   0.04725  -0.0223   0.9368   1.0000
   3.250   0.0446   0.06031   0.05005  -0.0265   0.9267   1.0000
   3.500   0.0654   0.06123   0.05091  -0.0276   0.9130   1.0000
   3.750   0.0871   0.06263   0.05225  -0.0289   0.9005   1.0000
   4.000   0.1200   0.06544   0.05496  -0.0322   0.8919   1.0000
   4.250   0.1466   0.06716   0.05663  -0.0341   0.8782   1.0000
   4.500   0.1632   0.06824   0.05768  -0.0346   0.8650   1.0000
   4.750   0.1849   0.07009   0.05949  -0.0359   0.8541   1.0000
   5.000   0.2232   0.07346   0.06280  -0.0398   0.8435   1.0000
   5.250   0.2349   0.07415   0.06350  -0.0395   0.8297   1.0000
   5.500   0.2512   0.07573   0.06507  -0.0400   0.8177   1.0000
   5.750   0.2891   0.07944   0.06875  -0.0437   0.8089   1.0000
   6.000   0.3026   0.08042   0.06973  -0.0437   0.7946   1.0000
   6.250   0.3138   0.08177   0.07111  -0.0437   0.7822   1.0000
   6.500   0.3388   0.08455   0.07389  -0.0456   0.7730   1.0000
   6.750   0.3663   0.08705   0.07641  -0.0475   0.7598   1.0000
   7.000   0.3729   0.08818   0.07757  -0.0470   0.7469   1.0000
   7.250   0.3896   0.09048   0.07991  -0.0478   0.7367   1.0000
   7.500   0.4268   0.09412   0.08358  -0.0509   0.7253   1.0000
   7.750   0.4285   0.09496   0.08447  -0.0499   0.7120   1.0000
   8.000   0.4394   0.09709   0.08664  -0.0503   0.7016   1.0000
   8.250   0.4745   0.10085   0.09046  -0.0530   0.6912   1.0000
   8.500   0.4809   0.10213   0.09180  -0.0527   0.6777   1.0000
   8.750   0.4862   0.10415   0.09388  -0.0527   0.6674   1.0000
   9.000   0.5134   0.10753   0.09731  -0.0547   0.6575   1.0000
   9.250   0.5323   0.10985   0.09972  -0.0556   0.6438   1.0000
   9.500   0.5301   0.11161   0.10154  -0.0552   0.6339   1.0000
   9.750   0.5503   0.11475   0.10475  -0.0566   0.6243   1.0000
  10.000   0.5794   0.11795   0.10803  -0.0582   0.6108   1.0000
  10.250   0.5726   0.11943   0.10958  -0.0577   0.6005   1.0000
  10.500   0.5864   0.12254   0.11279  -0.0588   0.5924   1.0000
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