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

Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 500,000
Max Cl/Cd: 67.09 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1070-il-500000.txt
Download as CSV file: xf-eh1070-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Eh 1.0/7.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5147   0.08129   0.07931   0.0054   1.0000   0.0139
  -8.250  -0.5175   0.07628   0.07432   0.0031   1.0000   0.0139
  -8.000  -0.5225   0.07101   0.06907   0.0001   1.0000   0.0139
  -7.750  -0.5325   0.06540   0.06348  -0.0047   1.0000   0.0139
  -7.500  -0.5435   0.06066   0.05870  -0.0073   1.0000   0.0139
  -7.250  -0.5930   0.06510   0.06288  -0.0095   1.0000   0.0138
  -7.000  -0.5862   0.06002   0.05767  -0.0112   1.0000   0.0139
  -6.750  -0.5782   0.05504   0.05252  -0.0121   1.0000   0.0139
  -6.500  -0.5685   0.05035   0.04762  -0.0124   1.0000   0.0140
  -6.250  -0.5764   0.04034   0.03725  -0.0129   1.0000   0.0151
  -6.000  -0.5612   0.03849   0.03534  -0.0129   1.0000   0.0163
  -5.750  -0.5431   0.03651   0.03324  -0.0127   1.0000   0.0183
  -5.500  -0.5242   0.03318   0.02964  -0.0119   1.0000   0.0200
  -5.250  -0.4963   0.03258   0.02875  -0.0107   1.0000   0.0232
  -5.000  -0.4748   0.03059   0.02644  -0.0095   1.0000   0.0236
  -4.750  -0.4637   0.02341   0.01869  -0.0082   1.0000   0.0254
  -4.500  -0.4392   0.01853   0.01322  -0.0056   1.0000   0.0147
  -4.250  -0.4156   0.01597   0.01042  -0.0044   1.0000   0.0127
  -4.000  -0.3918   0.01395   0.00811  -0.0031   1.0000   0.0126
  -3.750  -0.3679   0.01266   0.00668  -0.0019   1.0000   0.0132
  -3.500  -0.3439   0.01188   0.00580  -0.0009   1.0000   0.0141
  -3.250  -0.3237   0.01032   0.00413   0.0007   1.0000   0.0167
  -3.000  -0.3008   0.00972   0.00349   0.0017   1.0000   0.0186
  -2.750  -0.2779   0.00920   0.00290   0.0029   1.0000   0.0215
  -2.500  -0.2518   0.00847   0.00223   0.0033   0.9983   0.0523
  -2.250  -0.2149   0.00763   0.00193   0.0007   0.9897   0.1850
  -2.000  -0.1808   0.00640   0.00167  -0.0015   0.9797   0.4441
  -1.750  -0.1543   0.00518   0.00161  -0.0014   0.9634   0.7487
  -1.500  -0.1240   0.00494   0.00161  -0.0013   0.9437   0.8403
  -1.250  -0.0964   0.00487   0.00156  -0.0006   0.9180   0.8825
  -1.000  -0.0716   0.00490   0.00154   0.0008   0.8891   0.9160
  -0.750  -0.0452   0.00499   0.00152   0.0017   0.8595   0.9378
  -0.500  -0.0166   0.00510   0.00149   0.0020   0.8299   0.9540
  -0.250   0.0171   0.00524   0.00147   0.0011   0.8006   0.9670
   0.000   0.0593   0.00539   0.00147  -0.0017   0.7714   0.9760
   0.250   0.1102   0.00559   0.00150  -0.0064   0.7414   0.9863
   0.500   0.1532   0.00573   0.00149  -0.0097   0.7121   0.9935
   0.750   0.1913   0.00580   0.00143  -0.0120   0.6837   0.9976
   1.000   0.2234   0.00587   0.00137  -0.0130   0.6560   1.0000
   1.250   0.2463   0.00592   0.00133  -0.0121   0.6306   1.0000
   1.500   0.2696   0.00599   0.00131  -0.0113   0.6077   1.0000
   1.750   0.2932   0.00606   0.00131  -0.0105   0.5865   1.0000
   2.000   0.3169   0.00615   0.00135  -0.0098   0.5647   1.0000
   2.250   0.3407   0.00625   0.00139  -0.0090   0.5433   1.0000
   2.500   0.3646   0.00638   0.00145  -0.0083   0.5205   1.0000
   2.750   0.3888   0.00650   0.00152  -0.0076   0.4979   1.0000
   3.000   0.4130   0.00664   0.00161  -0.0069   0.4754   1.0000
   3.250   0.4373   0.00680   0.00174  -0.0062   0.4521   1.0000
   3.500   0.4617   0.00698   0.00187  -0.0056   0.4258   1.0000
   3.750   0.4857   0.00724   0.00199  -0.0049   0.3801   1.0000
   4.000   0.5090   0.00763   0.00215  -0.0042   0.3177   1.0000
   4.250   0.5326   0.00806   0.00237  -0.0036   0.2581   1.0000
   4.500   0.5526   0.00919   0.00285  -0.0028   0.1154   1.0000
   4.750   0.5723   0.01070   0.00394  -0.0015   0.0182   1.0000
   5.000   0.5963   0.01130   0.00466  -0.0007   0.0155   1.0000
   5.250   0.6199   0.01196   0.00540   0.0001   0.0133   1.0000
   5.500   0.6405   0.01328   0.00686   0.0013   0.0110   1.0000
   5.750   0.6609   0.01480   0.00852   0.0026   0.0102   1.0000
   6.000   0.6838   0.01590   0.00974   0.0036   0.0096   1.0000
   6.250   0.7056   0.01760   0.01162   0.0048   0.0094   1.0000
   6.500   0.7269   0.02001   0.01425   0.0061   0.0097   1.0000
   6.750   0.7495   0.02229   0.01667   0.0072   0.0112   1.0000
   7.250   0.7925   0.03000   0.02524   0.0109   0.0204   1.0000
   7.500   0.7889   0.03937   0.03521   0.0121   0.0186   1.0000
   7.750   0.8005   0.04233   0.03849   0.0132   0.0185   1.0000
   8.000   0.8232   0.04133   0.03777   0.0145   0.0171   1.0000
   8.250   0.8346   0.04408   0.04081   0.0156   0.0159   1.0000
   8.500   0.8391   0.04768   0.04468   0.0164   0.0151   1.0000
   8.750   0.8385   0.05150   0.04874   0.0170   0.0146   1.0000
   9.000   0.8316   0.05557   0.05301   0.0173   0.0142   1.0000
   9.250   0.8167   0.05930   0.05689   0.0178   0.0142   1.0000
   9.500   0.7954   0.06389   0.06160   0.0156   0.0144   1.0000
   9.750   0.7731   0.07118   0.06899   0.0089   0.0149   1.0000
  10.000   0.7536   0.08181   0.07968  -0.0002   0.0158   1.0000
<< Back to Eh 1.0/7.0 (eh1070-il)

Polar data table (+)

Polar graphs


<< Back to Eh 1.0/7.0 (eh1070-il)