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EPPLER 1210 AIRFOIL (e1210-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 1210 AIRFOIL (e1210-il)
Reynolds number: 50,000
Max Cl/Cd: 25.78 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e1210-il-50000-n5.txt
Download as CSV file: xf-e1210-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1210 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2051   0.10974   0.10304  -0.0369   1.0000   0.0870
  -9.500  -0.2105   0.10653   0.09994  -0.0379   1.0000   0.0877
  -9.250  -0.2111   0.10392   0.09748  -0.0379   1.0000   0.0891
  -9.000  -0.2085   0.10203   0.09573  -0.0370   1.0000   0.0909
  -8.750  -0.2166   0.09993   0.09382  -0.0363   0.9997   0.0914
  -8.500  -0.1992   0.09500   0.08890  -0.0425   0.9670   0.0932
  -8.250  -0.1881   0.08943   0.08331  -0.0500   0.9392   0.0956
  -8.000  -0.1831   0.08268   0.07651  -0.0591   0.9145   0.0969
  -7.750  -0.1471   0.08011   0.07387  -0.0625   0.8941   0.1010
  -7.500  -0.1371   0.07496   0.06863  -0.0690   0.8714   0.1034
  -7.250  -0.1389   0.06828   0.06181  -0.0776   0.8501   0.1054
  -7.000  -0.1426   0.06146   0.05477  -0.0850   0.8310   0.1077
  -6.750  -0.1233   0.05970   0.05290  -0.0855   0.8128   0.1107
  -6.500  -0.1196   0.05504   0.04798  -0.0891   0.7967   0.1139
  -6.250  -0.1250   0.04769   0.03991  -0.0947   0.7836   0.1180
  -6.000  -0.1022   0.04739   0.03966  -0.0934   0.7672   0.1209
  -5.750  -0.0860   0.04514   0.03713  -0.0939   0.7530   0.1249
  -5.500  -0.0722   0.04099   0.03220  -0.0959   0.7416   0.1302
  -5.250  -0.0504   0.04040   0.03166  -0.0949   0.7269   0.1335
  -5.000  -0.0286   0.03896   0.02996  -0.0947   0.7149   0.1379
  -4.750  -0.0078   0.03663   0.02697  -0.0952   0.7030   0.1438
  -4.500   0.0153   0.03593   0.02627  -0.0945   0.6906   0.1476
  -4.250   0.0402   0.03493   0.02504  -0.0941   0.6799   0.1529
  -4.000   0.0638   0.03357   0.02318  -0.0941   0.6680   0.1589
  -3.750   0.0900   0.03291   0.02248  -0.0936   0.6581   0.1636
  -3.500   0.1139   0.03224   0.02165  -0.0931   0.6461   0.1694
  -3.250   0.1416   0.03133   0.02034  -0.0929   0.6372   0.1761
  -3.000   0.1653   0.03092   0.01995  -0.0923   0.6256   0.1814
  -2.750   0.1930   0.03031   0.01907  -0.0919   0.6169   0.1888
  -2.500   0.2176   0.02989   0.01853  -0.0914   0.6061   0.1954
  -2.250   0.2446   0.02947   0.01802  -0.0909   0.5975   0.2027
  -2.000   0.2702   0.02920   0.01752  -0.0904   0.5875   0.2115
  -1.750   0.2967   0.02884   0.01718  -0.0899   0.5787   0.2192
  -1.500   0.3226   0.02866   0.01683  -0.0893   0.5701   0.2296
  -1.250   0.3479   0.02846   0.01664  -0.0886   0.5610   0.2390
  -1.000   0.3762   0.02824   0.01622  -0.0883   0.5538   0.2522
  -0.750   0.3998   0.02823   0.01630  -0.0876   0.5441   0.2647
  -0.500   0.4272   0.02803   0.01603  -0.0871   0.5371   0.2813
  -0.250   0.4513   0.02805   0.01612  -0.0864   0.5287   0.3005
   0.000   0.4764   0.02797   0.01611  -0.0858   0.5208   0.3263
   0.250   0.5039   0.02771   0.01589  -0.0853   0.5150   0.3636
   0.500   0.5241   0.02784   0.01637  -0.0843   0.5061   0.4155
   0.750   0.5467   0.02752   0.01645  -0.0830   0.4996   0.5115
   1.000   0.5646   0.02699   0.01656  -0.0800   0.4941   0.6929
   1.250   0.6011   0.02706   0.01690  -0.0812   0.4849   1.0000
   1.500   0.6279   0.02738   0.01690  -0.0808   0.4790   1.0000
   1.750   0.6513   0.02793   0.01725  -0.0802   0.4726   1.0000
   2.000   0.6727   0.02859   0.01779  -0.0794   0.4653   1.0000
   2.250   0.6992   0.02893   0.01789  -0.0789   0.4598   1.0000
   2.500   0.7219   0.02954   0.01836  -0.0782   0.4538   1.0000
   2.750   0.7411   0.03037   0.01914  -0.0773   0.4468   1.0000
   3.000   0.7667   0.03078   0.01938  -0.0767   0.4416   1.0000
   3.250   0.7925   0.03123   0.01964  -0.0763   0.4369   1.0000
   3.500   0.8062   0.03242   0.02089  -0.0749   0.4295   1.0000
   3.750   0.8293   0.03301   0.02137  -0.0742   0.4242   1.0000
   4.000   0.8578   0.03328   0.02145  -0.0740   0.4203   1.0000
   4.250   0.8684   0.03472   0.02297  -0.0724   0.4137   1.0000
   4.500   0.8852   0.03572   0.02396  -0.0713   0.4079   1.0000
   4.750   0.9110   0.03614   0.02426  -0.0708   0.4037   1.0000
   5.000   0.9354   0.03673   0.02473  -0.0703   0.3997   1.0000
   5.250   0.9336   0.03897   0.02716  -0.0680   0.3928   1.0000
   5.500   0.9508   0.03994   0.02810  -0.0670   0.3880   1.0000
   5.750   0.9791   0.04021   0.02825  -0.0667   0.3845   1.0000
   6.000   0.9834   0.04210   0.03019  -0.0648   0.3794   1.0000
   6.250   0.9674   0.04528   0.03355  -0.0620   0.3728   1.0000
   6.500   0.9842   0.04627   0.03451  -0.0610   0.3688   1.0000
   6.750   1.0169   0.04624   0.03437  -0.0609   0.3660   1.0000
   7.000   0.9019   0.05765   0.04616  -0.0564   0.3534   1.0000
   7.250   0.9201   0.05864   0.04712  -0.0557   0.3502   1.0000
   7.500   0.9512   0.05835   0.04674  -0.0549   0.3482   1.0000
   8.000   0.8790   0.07363   0.06224  -0.0575   0.3308   1.0000
   9.000   0.8087   0.09794   0.08676  -0.0636   0.3045   1.0000
   9.250   0.8145   0.10101   0.08984  -0.0640   0.3010   1.0000
   9.500   0.8300   0.10275   0.09156  -0.0638   0.2985   1.0000
   9.750   0.8508   0.10374   0.09252  -0.0632   0.2966   1.0000
  10.000   0.8149   0.11254   0.10143  -0.0666   0.2888   1.0000
  10.250   0.8196   0.11582   0.10473  -0.0673   0.2854   1.0000
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