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

EPPLER E1212 AIRFOIL (e1212-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER E1212 AIRFOIL (e1212-il)
Reynolds number: 100,000
Max Cl/Cd: 32.08 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1212-il-100000.txt
Download as CSV file: xf-e1212-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER E1212 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.3620   0.14384   0.13831  -0.0144   1.0000   0.1568
 -13.000  -0.3338   0.14060   0.13506  -0.0129   1.0000   0.1626
 -12.750  -0.3587   0.13889   0.13339  -0.0167   1.0000   0.1687
 -12.500  -0.3328   0.13391   0.12840  -0.0157   1.0000   0.1718
 -12.250  -0.3188   0.13129   0.12579  -0.0156   1.0000   0.1785
 -12.000  -0.3597   0.12962   0.12419  -0.0206   1.0000   0.1831
 -11.750  -0.3141   0.12462   0.11916  -0.0177   1.0000   0.1874
 -11.500  -0.3086   0.12228   0.11684  -0.0183   1.0000   0.1954
 -11.250  -0.3594   0.12017   0.11481  -0.0240   1.0000   0.1985
 -11.000  -0.2979   0.11565   0.11024  -0.0196   1.0000   0.2037
 -10.750  -0.3008   0.11356   0.10818  -0.0207   1.0000   0.2124
 -10.500  -0.3560   0.11093   0.10566  -0.0264   1.0000   0.2148
 -10.250  -0.2847   0.10695   0.10161  -0.0212   1.0000   0.2203
 -10.000  -0.2919   0.10505   0.09976  -0.0224   1.0000   0.2298
  -9.750  -0.3265   0.10146   0.09627  -0.0258   1.0000   0.2327
  -9.500  -0.2727   0.09867   0.09344  -0.0223   1.0000   0.2379
  -9.250  -0.2748   0.09671   0.09152  -0.0228   1.0000   0.2469
  -9.000  -0.3489   0.09388   0.08889  -0.0278   1.0000   0.2501
  -8.750  -0.2643   0.09085   0.08577  -0.0228   1.0000   0.2560
  -8.500  -0.2628   0.08900   0.08398  -0.0226   1.0000   0.2646
  -8.250  -0.3714   0.08173   0.07702  -0.0281   1.0000   0.2505
  -8.000  -0.3655   0.07948   0.07489  -0.0256   1.0000   0.2512
  -7.750  -0.3393   0.07645   0.07188  -0.0277   0.9887   0.2525
  -7.500  -0.4428   0.04467   0.03924  -0.0632   0.9387   0.2284
  -7.250  -0.4088   0.04129   0.03562  -0.0680   0.9216   0.2356
  -7.000  -0.3623   0.04084   0.03515  -0.0705   0.9041   0.2419
  -6.750  -0.3542   0.03562   0.02913  -0.0740   0.8827   0.2506
  -6.500  -0.3154   0.03565   0.02933  -0.0742   0.8628   0.2548
  -6.250  -0.2885   0.03499   0.02855  -0.0739   0.8425   0.2601
  -6.000  -0.2744   0.03287   0.02597  -0.0737   0.8232   0.2669
  -5.750  -0.2549   0.03155   0.02442  -0.0729   0.8050   0.2717
  -5.500  -0.2292   0.03129   0.02418  -0.0719   0.7876   0.2760
  -5.250  -0.2071   0.03054   0.02323  -0.0710   0.7720   0.2816
  -5.000  -0.1878   0.02931   0.02149  -0.0704   0.7578   0.2880
  -4.750  -0.1650   0.02852   0.02069  -0.0695   0.7419   0.2922
  -4.500  -0.1402   0.02817   0.02032  -0.0686   0.7278   0.2970
  -4.250  -0.1159   0.02757   0.01946  -0.0679   0.7155   0.3029
  -4.000  -0.0938   0.02688   0.01843  -0.0673   0.7011   0.3087
  -3.750  -0.0687   0.02628   0.01784  -0.0666   0.6892   0.3133
  -3.500  -0.0433   0.02594   0.01743  -0.0658   0.6766   0.3186
  -3.250  -0.0185   0.02553   0.01684  -0.0653   0.6644   0.3245
  -3.000   0.0072   0.02502   0.01604  -0.0648   0.6537   0.3303
  -2.750   0.0322   0.02470   0.01578  -0.0641   0.6415   0.3354
  -2.500   0.0591   0.02440   0.01532  -0.0636   0.6318   0.3415
  -2.250   0.0842   0.02418   0.01496  -0.0631   0.6197   0.3480
  -2.000   0.1113   0.02379   0.01444  -0.0626   0.6108   0.3538
  -1.750   0.1363   0.02364   0.01435  -0.0620   0.5991   0.3599
  -1.500   0.1640   0.02343   0.01392  -0.0616   0.5905   0.3673
  -1.250   0.1892   0.02326   0.01376  -0.0611   0.5795   0.3740
  -1.000   0.2165   0.02307   0.01352  -0.0606   0.5710   0.3813
  -0.750   0.2423   0.02306   0.01344  -0.0601   0.5609   0.3894
  -0.500   0.2692   0.02284   0.01320  -0.0597   0.5521   0.3973
  -0.250   0.2955   0.02285   0.01320  -0.0592   0.5436   0.4068
   0.000   0.3217   0.02277   0.01311  -0.0587   0.5344   0.4167
   0.250   0.3500   0.02269   0.01294  -0.0583   0.5275   0.4277
   0.500   0.3739   0.02274   0.01313  -0.0577   0.5175   0.4395
   0.750   0.4017   0.02268   0.01299  -0.0573   0.5104   0.4549
   1.000   0.4263   0.02277   0.01323  -0.0566   0.5022   0.4704
   1.250   0.4520   0.02273   0.01329  -0.0559   0.4941   0.4895
   1.500   0.4804   0.02266   0.01318  -0.0556   0.4882   0.5157
   1.750   0.5020   0.02282   0.01367  -0.0545   0.4793   0.5463
   2.000   0.5271   0.02268   0.01372  -0.0536   0.4726   0.5918
   2.250   0.5511   0.02259   0.01389  -0.0522   0.4666   0.6633
   2.500   0.5677   0.02251   0.01442  -0.0494   0.4588   0.7775
   2.750   0.6463   0.02226   0.01423  -0.0574   0.4501   0.9977
   3.000   0.6652   0.02289   0.01482  -0.0565   0.4428   1.0000
   3.250   0.6875   0.02329   0.01510  -0.0556   0.4361   1.0000
   3.500   0.7148   0.02355   0.01512  -0.0553   0.4310   1.0000
   3.750   0.7338   0.02430   0.01591  -0.0541   0.4244   1.0000
   4.000   0.7558   0.02483   0.01640  -0.0532   0.4179   1.0000
   4.250   0.7832   0.02510   0.01648  -0.0529   0.4128   1.0000
   4.500   0.8043   0.02583   0.01722  -0.0519   0.4070   1.0000
   4.750   0.8239   0.02655   0.01798  -0.0509   0.4005   1.0000
   5.000   0.8499   0.02694   0.01827  -0.0504   0.3956   1.0000
   5.250   0.8788   0.02739   0.01852  -0.0503   0.3915   1.0000
   5.500   0.8905   0.02858   0.01995  -0.0485   0.3847   1.0000
   5.750   0.9133   0.02914   0.02049  -0.0478   0.3793   1.0000
   6.000   0.9422   0.02942   0.02062  -0.0476   0.3751   1.0000
   6.250   0.9594   0.03056   0.02183  -0.0465   0.3703   1.0000
   6.500   0.9717   0.03188   0.02331  -0.0449   0.3647   1.0000
   6.750   0.9951   0.03247   0.02387  -0.0442   0.3601   1.0000
   7.000   1.0260   0.03270   0.02395  -0.0443   0.3564   1.0000
   7.250   1.0345   0.03435   0.02577  -0.0424   0.3515   1.0000
   7.500   1.0401   0.03610   0.02770  -0.0404   0.3462   1.0000
   7.750   1.0611   0.03691   0.02851  -0.0397   0.3423   1.0000
   8.000   1.0915   0.03721   0.02870  -0.0397   0.3391   1.0000
   8.250   1.1047   0.03875   0.03031  -0.0384   0.3354   1.0000
   8.500   1.0599   0.04359   0.03559  -0.0331   0.3296   1.0000
   8.750   1.0541   0.04597   0.03806  -0.0307   0.3254   1.0000
   9.000   1.1010   0.04515   0.03712  -0.0316   0.3226   1.0000
   9.250   1.1481   0.04465   0.03647  -0.0327   0.3202   1.0000
  10.000   0.5609   0.12501   0.11793  -0.0541   0.3458   1.0000
  10.250   0.5781   0.12733   0.12024  -0.0542   0.3416   1.0000
<< Back to EPPLER E1212 AIRFOIL (e1212-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER E1212 AIRFOIL (e1212-il)