Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

EPPLER E853 AIRFOIL (e853-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E853 AIRFOIL (e853-il)
Reynolds number: 100,000
Max Cl/Cd: 60.59 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e853-il-100000.txt
Download as CSV file: xf-e853-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E853 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3852   0.10212   0.09740  -0.0426   1.0000   0.1041
  -9.000  -0.4054   0.09923   0.09466  -0.0455   1.0000   0.1094
  -8.750  -0.4339   0.09648   0.09208  -0.0485   1.0000   0.1102
  -8.500  -0.4024   0.09260   0.08815  -0.0434   1.0000   0.1146
  -8.250  -0.4059   0.09003   0.08565  -0.0424   1.0000   0.1189
  -8.000  -0.4370   0.08784   0.08366  -0.0429   1.0000   0.1227
  -7.750  -0.4678   0.08616   0.08215  -0.0413   1.0000   0.1232
  -7.500  -0.5072   0.08280   0.07892  -0.0443   1.0000   0.1237
  -7.250  -0.4796   0.08153   0.07768  -0.0347   1.0000   0.1286
  -7.000  -0.4949   0.07960   0.07584  -0.0328   1.0000   0.1311
  -5.750  -0.4873   0.04317   0.03715  -0.0602   0.9904   0.0552
  -5.500  -0.4532   0.03739   0.03082  -0.0637   0.9862   0.0496
  -5.250  -0.4108   0.03279   0.02513  -0.0667   0.9828   0.0456
  -5.000  -0.3767   0.02991   0.02178  -0.0683   0.9777   0.0465
  -4.750  -0.3400   0.02810   0.01975  -0.0704   0.9728   0.0528
  -4.500  -0.3013   0.02592   0.01739  -0.0724   0.9692   0.0615
  -4.250  -0.2700   0.02441   0.01587  -0.0734   0.9630   0.0785
  -4.000  -0.2327   0.02276   0.01438  -0.0756   0.9588   0.1129
  -3.750  -0.1968   0.01993   0.01280  -0.0789   0.9551   0.2835
  -3.500  -0.1684   0.02014   0.01370  -0.0787   0.9480   0.5285
  -3.250  -0.1369   0.02081   0.01422  -0.0787   0.9420   0.5814
  -3.000  -0.1116   0.02126   0.01464  -0.0776   0.9348   0.6129
  -2.750  -0.0781   0.02161   0.01488  -0.0782   0.9301   0.6407
  -2.500  -0.0562   0.02186   0.01506  -0.0769   0.9222   0.6632
  -2.250  -0.0241   0.02211   0.01526  -0.0770   0.9178   0.6880
  -2.000  -0.0051   0.02232   0.01538  -0.0754   0.9101   0.7078
  -1.750   0.0277   0.02235   0.01533  -0.0762   0.9054   0.7239
  -1.500   0.0522   0.02245   0.01533  -0.0760   0.8989   0.7367
  -1.250   0.0833   0.02247   0.01525  -0.0771   0.8934   0.7491
  -1.000   0.1219   0.02240   0.01506  -0.0793   0.8900   0.7611
  -0.750   0.1373   0.02263   0.01527  -0.0776   0.8819   0.7717
  -0.500   0.1722   0.02259   0.01518  -0.0792   0.8778   0.7838
  -0.250   0.1924   0.02281   0.01538  -0.0784   0.8710   0.7958
   0.000   0.2220   0.02287   0.01539  -0.0790   0.8659   0.8092
   0.250   0.2594   0.02278   0.01529  -0.0809   0.8627   0.8234
   0.500   0.2697   0.02321   0.01575  -0.0785   0.8542   0.8389
   0.750   0.3037   0.02316   0.01571  -0.0796   0.8504   0.8568
   1.000   0.3162   0.02353   0.01613  -0.0774   0.8429   0.8777
   1.250   0.3437   0.02351   0.01617  -0.0774   0.8380   0.9047
   1.500   0.3898   0.02328   0.01603  -0.0808   0.8353   0.9442
   1.750   0.4203   0.02391   0.01668  -0.0838   0.8266   1.0000
   2.000   0.4696   0.02395   0.01668  -0.0887   0.8232   1.0000
   2.250   0.4944   0.02467   0.01735  -0.0899   0.8153   1.0000
   2.500   0.5352   0.02480   0.01747  -0.0928   0.8105   1.0000
   2.750   0.5860   0.02458   0.01731  -0.0969   0.8079   1.0000
   3.000   0.5977   0.02554   0.01827  -0.0952   0.7974   1.0000
   3.250   0.6494   0.02519   0.01799  -0.0991   0.7944   1.0000
   3.500   0.6631   0.02602   0.01889  -0.0975   0.7838   1.0000
   3.750   0.7165   0.02544   0.01842  -0.1011   0.7805   1.0000
   4.000   0.7335   0.02613   0.01918  -0.0998   0.7696   1.0000
   4.250   0.7637   0.02624   0.01941  -0.1000   0.7606   1.0000
   4.500   0.8176   0.02510   0.01847  -0.1028   0.7541   1.0000
   4.750   0.8532   0.02453   0.01805  -0.1030   0.7420   1.0000
   5.000   0.8995   0.02312   0.01680  -0.1039   0.7284   1.0000
   5.250   0.9436   0.02175   0.01557  -0.1044   0.7131   1.0000
   5.500   0.9823   0.02072   0.01473  -0.1044   0.6969   1.0000
   5.750   1.0066   0.02026   0.01444  -0.1025   0.6762   1.0000
   6.000   1.0371   0.01935   0.01366  -0.1012   0.6521   1.0000
   6.250   1.0602   0.01865   0.01308  -0.0987   0.6204   1.0000
   6.500   1.0777   0.01811   0.01262  -0.0953   0.5764   1.0000
   6.750   1.0870   0.01794   0.01233  -0.0906   0.4948   1.0000
   7.000   1.0775   0.01939   0.01259  -0.0835   0.3261   1.0000
   7.250   1.0627   0.02189   0.01412  -0.0770   0.2110   1.0000
   7.500   1.0549   0.02411   0.01578  -0.0717   0.1544   1.0000
   7.750   1.0535   0.02620   0.01754  -0.0675   0.1217   1.0000
   8.000   1.0589   0.02817   0.01934  -0.0644   0.0970   1.0000
   8.250   1.0701   0.03016   0.02126  -0.0621   0.0777   1.0000
   8.500   1.0895   0.03247   0.02350  -0.0608   0.0626   1.0000
   8.750   1.1145   0.03497   0.02601  -0.0604   0.0505   1.0000
   9.000   1.1507   0.03841   0.02957  -0.0616   0.0419   1.0000
   9.250   1.1756   0.04114   0.03249  -0.0613   0.0369   1.0000
   9.500   1.2094   0.04668   0.03830  -0.0626   0.0348   1.0000
   9.750   1.2247   0.05203   0.04413  -0.0613   0.0344   1.0000
  10.000   1.2283   0.05635   0.04886  -0.0586   0.0341   1.0000
  10.250   1.2237   0.05884   0.05179  -0.0548   0.0337   1.0000
  10.500   1.2170   0.06212   0.05543  -0.0512   0.0336   1.0000
  10.750   1.2047   0.06491   0.05854  -0.0472   0.0333   1.0000
  11.000   1.1917   0.06851   0.06243  -0.0440   0.0334   1.0000
  11.250   1.1765   0.07228   0.06646  -0.0414   0.0334   1.0000
  11.500   1.1596   0.07634   0.07075  -0.0395   0.0335   1.0000
  11.750   1.1438   0.08136   0.07597  -0.0385   0.0339   1.0000
  12.000   0.9923   0.08556   0.08125  -0.0328   0.0397   1.0000
  12.250   0.9605   0.09225   0.08815  -0.0348   0.0407   1.0000
<< Back to EPPLER E853 AIRFOIL (e853-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E853 AIRFOIL (e853-il)