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EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER EA 8(-1)-006 AIRFOIL (ea81006-il)
Reynolds number: 100,000
Max Cl/Cd: 23.38 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ea81006-il-100000-n5.txt
Download as CSV file: xf-ea81006-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER EA 8(-1)-006 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6869   0.08406   0.07922  -0.0044   1.0000   0.0221
  -9.000  -0.6946   0.07815   0.07336  -0.0092   1.0000   0.0219
  -8.750  -0.7048   0.07282   0.06801  -0.0121   1.0000   0.0218
  -8.500  -0.7135   0.06722   0.06232  -0.0142   1.0000   0.0217
  -8.250  -0.7203   0.06143   0.05637  -0.0156   1.0000   0.0217
  -8.000  -0.7245   0.05573   0.05040  -0.0162   1.0000   0.0217
  -7.750  -0.7265   0.04977   0.04404  -0.0159   1.0000   0.0219
  -7.500  -0.7163   0.04694   0.04100  -0.0153   1.0000   0.0230
  -7.250  -0.7064   0.04338   0.03706  -0.0145   1.0000   0.0250
  -7.000  -0.6947   0.03990   0.03306  -0.0134   1.0000   0.0270
  -6.750  -0.6810   0.03632   0.02896  -0.0120   1.0000   0.0287
  -6.500  -0.6649   0.03312   0.02524  -0.0106   1.0000   0.0308
  -6.250  -0.6462   0.03073   0.02234  -0.0093   1.0000   0.0345
  -6.000  -0.6256   0.02950   0.02081  -0.0083   1.0000   0.0401
  -5.750  -0.6047   0.02753   0.01821  -0.0069   1.0000   0.0454
  -5.500  -0.5834   0.02603   0.01664  -0.0063   1.0000   0.0514
  -5.250  -0.5611   0.02455   0.01477  -0.0053   1.0000   0.0575
  -5.000  -0.5376   0.02327   0.01317  -0.0043   1.0000   0.0618
  -4.750  -0.5141   0.02199   0.01183  -0.0037   1.0000   0.0664
  -4.500  -0.4897   0.02095   0.01053  -0.0028   1.0000   0.0683
  -4.250  -0.4655   0.01995   0.00945  -0.0020   1.0000   0.0699
  -4.000  -0.4423   0.01897   0.00853  -0.0012   1.0000   0.0726
  -3.750  -0.4191   0.01822   0.00774  -0.0003   1.0000   0.0744
  -3.500  -0.3963   0.01755   0.00704   0.0008   1.0000   0.0756
  -3.250  -0.3746   0.01683   0.00633   0.0019   1.0000   0.0771
  -3.000  -0.3530   0.01622   0.00575   0.0030   1.0000   0.0794
  -2.750  -0.3309   0.01574   0.00524   0.0041   1.0000   0.0816
  -2.500  -0.3085   0.01534   0.00479   0.0052   1.0000   0.0831
  -2.250  -0.2864   0.01490   0.00435   0.0063   1.0000   0.0851
  -2.000  -0.2641   0.01453   0.00396   0.0073   1.0000   0.0887
  -1.750  -0.2416   0.01423   0.00366   0.0083   1.0000   0.0931
  -1.500  -0.2192   0.01391   0.00341   0.0094   1.0000   0.0991
  -1.250  -0.1976   0.01343   0.00324   0.0104   1.0000   0.1328
  -1.000  -0.1768   0.01267   0.00311   0.0114   1.0000   0.2887
  -0.750  -0.1612   0.01125   0.00311   0.0135   1.0000   0.6051
  -0.500  -0.1327   0.01078   0.00354   0.0150   1.0000   0.8629
  -0.250  -0.0644   0.01088   0.00374   0.0075   1.0000   0.9635
   0.250   0.0642   0.01088   0.00374  -0.0075   0.9638   1.0000
   0.500   0.1330   0.01078   0.00353  -0.0150   0.8620   1.0000
   0.750   0.1613   0.01123   0.00312  -0.0135   0.6087   1.0000
   1.000   0.1769   0.01266   0.00311  -0.0113   0.2909   1.0000
   1.250   0.1975   0.01344   0.00324  -0.0104   0.1307   1.0000
   1.500   0.2192   0.01391   0.00341  -0.0093   0.0991   1.0000
   1.750   0.2416   0.01423   0.00366  -0.0083   0.0931   1.0000
   2.000   0.2641   0.01453   0.00396  -0.0073   0.0887   1.0000
   2.250   0.2864   0.01490   0.00435  -0.0062   0.0851   1.0000
   2.500   0.3085   0.01534   0.00479  -0.0052   0.0831   1.0000
   2.750   0.3309   0.01574   0.00524  -0.0041   0.0815   1.0000
   3.000   0.3530   0.01621   0.00574  -0.0030   0.0795   1.0000
   3.250   0.3746   0.01683   0.00633  -0.0019   0.0771   1.0000
   3.500   0.3963   0.01755   0.00703  -0.0008   0.0755   1.0000
   3.750   0.4191   0.01821   0.00773   0.0003   0.0745   1.0000
   4.000   0.4423   0.01897   0.00853   0.0012   0.0727   1.0000
   4.250   0.4655   0.01994   0.00944   0.0020   0.0700   1.0000
   4.500   0.4896   0.02094   0.01053   0.0029   0.0683   1.0000
   4.750   0.5140   0.02199   0.01183   0.0037   0.0665   1.0000
   5.000   0.5375   0.02326   0.01317   0.0044   0.0618   1.0000
   5.250   0.5610   0.02455   0.01477   0.0053   0.0575   1.0000
   5.500   0.5833   0.02602   0.01663   0.0063   0.0514   1.0000
   5.750   0.6046   0.02754   0.01822   0.0069   0.0454   1.0000
   6.000   0.6256   0.02950   0.02082   0.0083   0.0401   1.0000
   6.250   0.6461   0.03074   0.02235   0.0093   0.0345   1.0000
   6.500   0.6648   0.03310   0.02522   0.0106   0.0308   1.0000
   6.750   0.6809   0.03633   0.02898   0.0121   0.0287   1.0000
   7.000   0.6946   0.03990   0.03306   0.0134   0.0270   1.0000
   7.250   0.7063   0.04339   0.03708   0.0145   0.0249   1.0000
   7.500   0.7161   0.04699   0.04106   0.0154   0.0230   1.0000
   7.750   0.7262   0.04982   0.04410   0.0159   0.0219   1.0000
   8.000   0.7247   0.05564   0.05031   0.0162   0.0217   1.0000
   8.250   0.7199   0.06155   0.05649   0.0156   0.0216   1.0000
   8.500   0.7138   0.06715   0.06225   0.0142   0.0217   1.0000
   8.750   0.7045   0.07294   0.06813   0.0120   0.0218   1.0000
   9.000   0.6947   0.07819   0.07339   0.0091   0.0219   1.0000
   9.250   0.6871   0.08410   0.07926   0.0043   0.0221   1.0000
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