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

EPPLER 1213 AIRFOIL (e1213-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 1213 AIRFOIL (e1213-il)
Reynolds number: 100,000
Max Cl/Cd: 30.63 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e1213-il-100000.txt
Download as CSV file: xf-e1213-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 1213 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
  ------ -------- --------- --------- -------- -------- --------
 -14.750  -0.5288   0.15050   0.14485   0.0043   1.0000   0.1036
 -14.500  -0.8540   0.07622   0.07027  -0.0382   1.0000   0.0947
 -14.250  -0.9095   0.06464   0.05842  -0.0460   1.0000   0.0948
 -14.000  -0.8783   0.06634   0.06038  -0.0436   1.0000   0.1022
 -13.750  -0.9302   0.05611   0.04985  -0.0506   1.0000   0.1026
 -13.500  -0.9092   0.05629   0.05029  -0.0493   1.0000   0.1103
 -13.250  -0.9500   0.04881   0.04254  -0.0539   1.0000   0.1120
 -13.000  -0.9850   0.04343   0.03684  -0.0557   1.0000   0.1143
 -12.750  -0.9714   0.04318   0.03687  -0.0547   1.0000   0.1228
 -12.500  -0.9703   0.04207   0.03584  -0.0534   1.0000   0.1297
 -12.250  -0.9855   0.04014   0.03378  -0.0508   1.0000   0.1353
 -12.000  -0.9636   0.04103   0.03488  -0.0492   1.0000   0.1439
 -11.750  -0.4079   0.12289   0.11734   0.0003   1.0000   0.2161
 -11.500  -0.4460   0.11990   0.11440  -0.0042   1.0000   0.2234
 -11.250  -0.4696   0.10974   0.10422  -0.0096   1.0000   0.2074
 -11.000  -0.4619   0.10507   0.09952  -0.0105   1.0000   0.2068
 -10.750  -0.4623   0.10037   0.09483  -0.0119   1.0000   0.2091
 -10.500  -0.4656   0.09535   0.08981  -0.0137   1.0000   0.2094
 -10.250  -0.5038   0.08564   0.08011  -0.0195   1.0000   0.2068
 -10.000  -0.7540   0.05074   0.04513  -0.0377   1.0000   0.2057
  -9.250  -0.7539   0.04422   0.03827  -0.0335   1.0000   0.2224
  -9.000  -0.7319   0.04379   0.03790  -0.0324   1.0000   0.2271
  -8.750  -0.6902   0.04561   0.03992  -0.0314   1.0000   0.2320
  -8.500  -0.7147   0.04098   0.03494  -0.0292   1.0000   0.2385
  -8.250  -0.7141   0.03884   0.03267  -0.0270   1.0000   0.2434
  -8.000  -0.6868   0.03960   0.03368  -0.0256   1.0000   0.2473
  -7.750  -0.6970   0.03889   0.03300  -0.0213   1.0000   0.2506
  -7.500  -0.6782   0.03550   0.02904  -0.0251   0.9835   0.2596
  -7.250  -0.6165   0.03574   0.02954  -0.0300   0.9689   0.2657
  -7.000  -0.5723   0.03345   0.02682  -0.0354   0.9504   0.2754
  -6.750  -0.5253   0.03231   0.02568  -0.0390   0.9305   0.2816
  -6.500  -0.4819   0.03155   0.02485  -0.0417   0.9083   0.2884
  -6.250  -0.4575   0.02986   0.02261  -0.0425   0.8842   0.2960
  -6.000  -0.4266   0.02926   0.02206  -0.0423   0.8620   0.3005
  -5.750  -0.4018   0.02886   0.02157  -0.0413   0.8390   0.3057
  -5.500  -0.3820   0.02805   0.02045  -0.0401   0.8182   0.3120
  -5.250  -0.3610   0.02722   0.01938  -0.0390   0.7994   0.3170
  -5.000  -0.3365   0.02687   0.01900  -0.0378   0.7815   0.3217
  -4.750  -0.3134   0.02643   0.01840  -0.0367   0.7649   0.3275
  -4.500  -0.2922   0.02583   0.01746  -0.0357   0.7483   0.3336
  -4.250  -0.2678   0.02527   0.01690  -0.0348   0.7323   0.3381
  -4.000  -0.2429   0.02497   0.01653  -0.0339   0.7177   0.3437
  -3.750  -0.2187   0.02458   0.01589  -0.0329   0.7051   0.3500
  -3.500  -0.1946   0.02409   0.01526  -0.0323   0.6901   0.3555
  -3.250  -0.1689   0.02379   0.01497  -0.0316   0.6766   0.3609
  -3.000  -0.1433   0.02352   0.01451  -0.0307   0.6654   0.3675
  -2.750  -0.1184   0.02322   0.01404  -0.0302   0.6515   0.3739
  -2.500  -0.0922   0.02289   0.01375  -0.0296   0.6397   0.3795
  -2.250  -0.0662   0.02268   0.01344  -0.0289   0.6281   0.3866
  -2.000  -0.0405   0.02250   0.01311  -0.0284   0.6161   0.3940
  -1.750  -0.0136   0.02221   0.01281  -0.0277   0.6061   0.4004
  -1.500   0.0122   0.02210   0.01268  -0.0273   0.5939   0.4084
  -1.250   0.0392   0.02189   0.01232  -0.0267   0.5850   0.4166
  -1.000   0.0650   0.02178   0.01232  -0.0262   0.5728   0.4252
  -0.750   0.0921   0.02166   0.01202  -0.0257   0.5642   0.4348
  -0.500   0.1180   0.02154   0.01205  -0.0252   0.5527   0.4444
  -0.250   0.1451   0.02145   0.01184  -0.0246   0.5444   0.4563
   0.000   0.1707   0.02140   0.01191  -0.0241   0.5334   0.4688
   0.250   0.1980   0.02128   0.01173  -0.0235   0.5259   0.4826
   0.500   0.2230   0.02127   0.01189  -0.0230   0.5149   0.4988
   0.750   0.2501   0.02115   0.01174  -0.0223   0.5076   0.5190
   1.000   0.2743   0.02118   0.01201  -0.0216   0.4973   0.5421
   1.250   0.3007   0.02104   0.01192  -0.0208   0.4897   0.5733
   1.500   0.3245   0.02104   0.01221  -0.0199   0.4807   0.6120
   1.750   0.3494   0.02085   0.01228  -0.0187   0.4726   0.6679
   2.000   0.3736   0.02074   0.01250  -0.0171   0.4653   0.7483
   2.250   0.4097   0.02073   0.01287  -0.0173   0.4554   0.8586
   2.500   0.4797   0.02107   0.01318  -0.0242   0.4445   0.9411
   2.750   0.5500   0.02132   0.01326  -0.0318   0.4335   0.9862
   3.000   0.5907   0.02170   0.01357  -0.0349   0.4244   1.0000
   3.250   0.6080   0.02199   0.01377  -0.0334   0.4175   1.0000
   3.500   0.6302   0.02225   0.01378  -0.0323   0.4124   1.0000
   3.750   0.6463   0.02290   0.01453  -0.0308   0.4045   1.0000
   4.000   0.6677   0.02322   0.01477  -0.0296   0.3981   1.0000
   4.250   0.6920   0.02356   0.01489  -0.0288   0.3929   1.0000
   4.500   0.7081   0.02431   0.01579  -0.0272   0.3851   1.0000
   4.750   0.7307   0.02469   0.01609  -0.0262   0.3793   1.0000
   5.000   0.7560   0.02510   0.01631  -0.0255   0.3747   1.0000
   5.250   0.7705   0.02603   0.01744  -0.0239   0.3674   1.0000
   5.500   0.7931   0.02645   0.01780  -0.0229   0.3617   1.0000
   5.750   0.8204   0.02678   0.01792  -0.0224   0.3573   1.0000
   6.000   0.8326   0.02797   0.01935  -0.0206   0.3507   1.0000
   6.250   0.8531   0.02866   0.02006  -0.0196   0.3453   1.0000
   6.500   0.8792   0.02902   0.02030  -0.0190   0.3411   1.0000
   6.750   0.8955   0.03016   0.02153  -0.0177   0.3360   1.0000
   7.000   0.9095   0.03132   0.02284  -0.0162   0.3302   1.0000
   7.250   0.9330   0.03183   0.02329  -0.0155   0.3257   1.0000
   7.500   0.9620   0.03220   0.02351  -0.0153   0.3222   1.0000
   7.750   0.9641   0.03435   0.02597  -0.0130   0.3172   1.0000
   8.000   0.9727   0.03600   0.02779  -0.0114   0.3124   1.0000
   8.250   0.9954   0.03662   0.02839  -0.0107   0.3086   1.0000
   8.500   1.0273   0.03677   0.02839  -0.0106   0.3053   1.0000
   8.750   1.0184   0.03970   0.03162  -0.0080   0.3009   1.0000
   9.000   0.9829   0.04439   0.03667  -0.0042   0.2963   1.0000
   9.500   1.0158   0.04662   0.03893  -0.0024   0.2901   1.0000
   9.750   1.0744   0.04524   0.03732  -0.0037   0.2875   1.0000
  10.000   0.5053   0.11800   0.11095  -0.0300   0.3273   1.0000
  10.250   0.5269   0.12052   0.11347  -0.0299   0.3244   1.0000
<< Back to EPPLER 1213 AIRFOIL (e1213-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 1213 AIRFOIL (e1213-il)