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Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 1,000,000
Max Cl/Cd: 76.98 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh1070-il-1000000.txt
Download as CSV file: xf-eh1070-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6080   0.08152   0.08003   0.0043   1.0000   0.0063
  -8.000  -0.6176   0.07690   0.07544   0.0005   1.0000   0.0061
  -7.750  -0.6219   0.07145   0.06997  -0.0039   1.0000   0.0066
  -7.500  -0.6215   0.06611   0.06457  -0.0073   1.0000   0.0064
  -7.250  -0.6143   0.06078   0.05916  -0.0098   1.0000   0.0070
  -7.000  -0.5978   0.05686   0.05515  -0.0110   1.0000   0.0075
  -6.750  -0.5884   0.05138   0.04952  -0.0120   1.0000   0.0076
  -6.500  -0.5778   0.04581   0.04377  -0.0123   1.0000   0.0076
  -6.250  -0.5658   0.04043   0.03815  -0.0120   1.0000   0.0077
  -6.000  -0.5523   0.03571   0.03315  -0.0112   1.0000   0.0077
  -5.750  -0.5531   0.02575   0.02249  -0.0094   1.0000   0.0085
  -5.500  -0.5331   0.02364   0.02017  -0.0087   1.0000   0.0090
  -5.250  -0.5109   0.02250   0.01892  -0.0083   1.0000   0.0098
  -5.000  -0.4875   0.02154   0.01783  -0.0077   1.0000   0.0114
  -4.750  -0.4630   0.02026   0.01638  -0.0068   1.0000   0.0128
  -4.500  -0.4425   0.01467   0.01004  -0.0039   1.0000   0.0081
  -4.250  -0.4196   0.01262   0.00780  -0.0025   1.0000   0.0072
  -4.000  -0.3968   0.01112   0.00611  -0.0012   1.0000   0.0072
  -3.750  -0.3737   0.01022   0.00511  -0.0001   1.0000   0.0075
  -3.500  -0.3498   0.00873   0.00344   0.0008   0.9986   0.0091
  -3.250  -0.3140   0.00820   0.00284  -0.0009   0.9918   0.0104
  -3.000  -0.2776   0.00784   0.00245  -0.0028   0.9832   0.0120
  -2.750  -0.2423   0.00751   0.00206  -0.0044   0.9694   0.0138
  -2.500  -0.2114   0.00693   0.00155  -0.0049   0.9457   0.0544
  -2.250  -0.1859   0.00665   0.00137  -0.0043   0.9136   0.1035
  -2.000  -0.1623   0.00632   0.00121  -0.0034   0.8799   0.1808
  -1.750  -0.1392   0.00583   0.00105  -0.0026   0.8484   0.3085
  -1.500  -0.1181   0.00496   0.00088  -0.0016   0.8184   0.5478
  -1.250  -0.0959   0.00446   0.00079  -0.0005   0.7895   0.7023
  -1.000  -0.0720   0.00427   0.00076   0.0005   0.7621   0.7831
  -0.750  -0.0473   0.00421   0.00075   0.0014   0.7362   0.8360
  -0.250   0.0033   0.00421   0.00073   0.0029   0.6869   0.8959
   0.000   0.0291   0.00426   0.00072   0.0036   0.6642   0.9163
   0.500   0.0816   0.00440   0.00076   0.0047   0.6204   0.9540
   0.750   0.1108   0.00451   0.00079   0.0046   0.5973   0.9693
   1.000   0.1465   0.00466   0.00085   0.0030   0.5744   0.9820
   1.250   0.1878   0.00477   0.00088   0.0001   0.5535   0.9854
   1.500   0.2240   0.00487   0.00091  -0.0018   0.5327   0.9883
   1.750   0.2586   0.00497   0.00093  -0.0034   0.5119   0.9914
   2.000   0.2920   0.00509   0.00099  -0.0047   0.4889   0.9943
   2.750   0.3935   0.00541   0.00111  -0.0090   0.4208   1.0000
   3.000   0.4173   0.00555   0.00119  -0.0083   0.3965   1.0000
   3.250   0.4411   0.00573   0.00127  -0.0076   0.3630   1.0000
   3.500   0.4644   0.00604   0.00138  -0.0068   0.3093   1.0000
   3.750   0.4878   0.00636   0.00152  -0.0061   0.2615   1.0000
   4.000   0.5105   0.00685   0.00173  -0.0054   0.1925   1.0000
   4.250   0.5321   0.00758   0.00207  -0.0046   0.1027   1.0000
   4.500   0.5530   0.00859   0.00269  -0.0035   0.0146   1.0000
   4.750   0.5770   0.00908   0.00327  -0.0027   0.0100   1.0000
   5.000   0.6014   0.00942   0.00366  -0.0020   0.0086   1.0000
   5.250   0.6256   0.00986   0.00417  -0.0013   0.0077   1.0000
   5.500   0.6492   0.01043   0.00480  -0.0005   0.0065   1.0000
   5.750   0.6682   0.01210   0.00670   0.0011   0.0054   1.0000
   6.000   0.6911   0.01296   0.00767   0.0020   0.0052   1.0000
   6.250   0.7125   0.01432   0.00917   0.0032   0.0051   1.0000
   6.500   0.7327   0.01639   0.01141   0.0045   0.0053   1.0000
   7.000   0.7627   0.02801   0.02403   0.0091   0.0104   1.0000
   7.250   0.7821   0.02969   0.02595   0.0101   0.0102   1.0000
   7.500   0.8110   0.02836   0.02466   0.0106   0.0088   1.0000
   7.750   0.8302   0.03000   0.02649   0.0114   0.0079   1.0000
   8.000   0.8465   0.03204   0.02872   0.0123   0.0073   1.0000
   8.250   0.8621   0.03367   0.03049   0.0129   0.0067   1.0000
   8.500   0.8746   0.03589   0.03286   0.0136   0.0065   1.0000
   8.750   0.8825   0.03867   0.03585   0.0146   0.0062   1.0000
   9.000   0.8841   0.04212   0.03955   0.0156   0.0061   1.0000
   9.250   0.8768   0.04631   0.04396   0.0166   0.0059   1.0000
   9.500   0.8557   0.05097   0.04884   0.0181   0.0058   1.0000
   9.750   0.8365   0.05429   0.05228   0.0187   0.0057   1.0000
  10.000   0.8244   0.05801   0.05611   0.0165   0.0058   1.0000
  10.250   0.8092   0.06371   0.06190   0.0115   0.0058   1.0000
  10.500   0.8046   0.06927   0.06757   0.0064   0.0060   1.0000
  10.750   0.7853   0.08013   0.07856  -0.0023   0.0062   1.0000
  11.000   0.7569   0.09806   0.09652  -0.0116   0.0065   1.0000
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