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Eh 1.0/7.0 (eh1070-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: Eh 1.0/7.0 (eh1070-il)
Reynolds number: 200,000
Max Cl/Cd: 49.86 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-eh1070-il-200000-n5.txt
Download as CSV file: xf-eh1070-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6036   0.08399   0.08071   0.0049   1.0000   0.0088
  -8.000  -0.6093   0.07914   0.07593   0.0011   1.0000   0.0088
  -7.750  -0.6143   0.07422   0.07100  -0.0028   1.0000   0.0085
  -7.500  -0.6147   0.06880   0.06554  -0.0064   1.0000   0.0085
  -7.250  -0.6126   0.06328   0.05993  -0.0092   1.0000   0.0084
  -7.000  -0.6074   0.05783   0.05434  -0.0112   1.0000   0.0085
  -6.750  -0.5989   0.05264   0.04896  -0.0124   1.0000   0.0089
  -6.500  -0.5863   0.04757   0.04364  -0.0128   1.0000   0.0100
  -6.250  -0.5702   0.04298   0.03875  -0.0126   1.0000   0.0110
  -6.000  -0.5554   0.03852   0.03394  -0.0120   1.0000   0.0112
  -5.750  -0.5408   0.03391   0.02891  -0.0111   1.0000   0.0110
  -5.500  -0.5236   0.02989   0.02442  -0.0100   1.0000   0.0109
  -5.250  -0.5039   0.02650   0.02054  -0.0089   1.0000   0.0109
  -5.000  -0.4823   0.02360   0.01709  -0.0077   1.0000   0.0109
  -4.750  -0.4591   0.02123   0.01431  -0.0066   1.0000   0.0111
  -4.500  -0.4349   0.01940   0.01214  -0.0057   1.0000   0.0115
  -4.250  -0.4117   0.01703   0.00947  -0.0045   1.0000   0.0123
  -4.000  -0.3885   0.01566   0.00800  -0.0037   1.0000   0.0148
  -3.750  -0.3643   0.01487   0.00714  -0.0030   1.0000   0.0174
  -3.500  -0.3410   0.01384   0.00600  -0.0019   1.0000   0.0182
  -3.250  -0.3177   0.01303   0.00508  -0.0009   1.0000   0.0193
  -3.000  -0.2942   0.01241   0.00433   0.0000   1.0000   0.0211
  -2.750  -0.2711   0.01173   0.00356   0.0010   1.0000   0.0295
  -2.500  -0.2485   0.01115   0.00313   0.0019   1.0000   0.0644
  -2.250  -0.2189   0.01002   0.00281   0.0006   0.9889   0.2309
  -2.000  -0.1901   0.00863   0.00253  -0.0005   0.9715   0.5070
  -1.750  -0.1640   0.00778   0.00253   0.0003   0.9519   0.7319
  -1.500  -0.1324   0.00760   0.00256   0.0004   0.9305   0.8347
  -1.250  -0.0962   0.00759   0.00248  -0.0005   0.9072   0.8840
  -1.000  -0.0588   0.00764   0.00242  -0.0018   0.8811   0.9174
  -0.750  -0.0169   0.00773   0.00239  -0.0040   0.8540   0.9474
  -0.500   0.0329   0.00785   0.00235  -0.0080   0.8251   0.9725
  -0.250   0.0716   0.00792   0.00222  -0.0101   0.7937   0.9827
   0.000   0.1072   0.00796   0.00209  -0.0117   0.7621   0.9890
   0.250   0.1411   0.00802   0.00199  -0.0129   0.7327   0.9953
   0.500   0.1732   0.00808   0.00190  -0.0139   0.7046   1.0000
   0.750   0.1957   0.00816   0.00186  -0.0129   0.6791   1.0000
   1.000   0.2186   0.00825   0.00183  -0.0119   0.6547   1.0000
   1.250   0.2417   0.00835   0.00182  -0.0110   0.6315   1.0000
   1.500   0.2652   0.00846   0.00184  -0.0101   0.6089   1.0000
   1.750   0.2888   0.00858   0.00188  -0.0093   0.5867   1.0000
   2.000   0.3125   0.00872   0.00198  -0.0085   0.5644   1.0000
   2.250   0.3365   0.00886   0.00206  -0.0077   0.5420   1.0000
   2.500   0.3606   0.00901   0.00216  -0.0070   0.5214   1.0000
   2.750   0.3849   0.00917   0.00229  -0.0063   0.5013   1.0000
   3.000   0.4093   0.00934   0.00244  -0.0057   0.4806   1.0000
   3.250   0.4336   0.00954   0.00260  -0.0050   0.4576   1.0000
   3.500   0.4580   0.00974   0.00285  -0.0043   0.4334   1.0000
   3.750   0.4825   0.00996   0.00307  -0.0037   0.4081   1.0000
   4.000   0.5068   0.01022   0.00331  -0.0030   0.3798   1.0000
   4.250   0.5300   0.01063   0.00354  -0.0023   0.3222   1.0000
   4.500   0.5505   0.01157   0.00386  -0.0016   0.1953   1.0000
   4.750   0.5711   0.01272   0.00448  -0.0009   0.0853   1.0000
   5.000   0.5918   0.01400   0.00541  -0.0001   0.0224   1.0000
   5.250   0.6149   0.01489   0.00651   0.0008   0.0163   1.0000
   5.500   0.6373   0.01591   0.00774   0.0018   0.0144   1.0000
   5.750   0.6592   0.01699   0.00900   0.0028   0.0134   1.0000
   6.000   0.6817   0.01790   0.01003   0.0036   0.0111   1.0000
   6.250   0.7037   0.01887   0.01107   0.0043   0.0090   1.0000
   6.500   0.7230   0.02070   0.01304   0.0054   0.0081   1.0000
   6.750   0.7410   0.02360   0.01618   0.0069   0.0075   1.0000
   7.000   0.7613   0.02585   0.01869   0.0080   0.0074   1.0000
   7.250   0.7810   0.02809   0.02127   0.0090   0.0072   1.0000
   7.500   0.7985   0.03075   0.02431   0.0102   0.0072   1.0000
   7.750   0.8133   0.03378   0.02775   0.0113   0.0071   1.0000
   8.000   0.8247   0.03713   0.03152   0.0125   0.0071   1.0000
   8.250   0.8323   0.04076   0.03563   0.0136   0.0072   1.0000
   8.500   0.8353   0.04467   0.03991   0.0147   0.0072   1.0000
   8.750   0.8331   0.04877   0.04434   0.0155   0.0073   1.0000
   9.000   0.8254   0.05298   0.04880   0.0161   0.0074   1.0000
   9.500   0.8018   0.06006   0.05622   0.0160   0.0075   1.0000
   9.750   0.7871   0.06490   0.06120   0.0126   0.0075   1.0000
  10.000   0.7787   0.06997   0.06639   0.0081   0.0076   1.0000
  10.250   0.7680   0.07671   0.07322   0.0023   0.0077   1.0000
  10.500   0.7600   0.08403   0.08062  -0.0034   0.0077   1.0000
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