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EPPLER 399 AIRFOIL (e399-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 399 AIRFOIL (e399-il)
Reynolds number: 50,000
Max Cl/Cd: 6.33 at α=8.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e399-il-50000.txt
Download as CSV file: xf-e399-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 399 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3823   0.12207   0.11665  -0.0156   1.0000   0.2727
  -7.750  -0.3885   0.12054   0.11516  -0.0136   1.0000   0.2808
  -7.500  -0.4207   0.12112   0.11588  -0.0119   1.0000   0.2844
  -7.250  -0.4025   0.11673   0.11145  -0.0100   1.0000   0.2946
  -7.000  -0.4281   0.11648   0.11131  -0.0080   1.0000   0.3002
  -6.750  -0.4318   0.11367   0.10854  -0.0063   1.0000   0.3055
  -6.500  -0.4353   0.11169   0.10657  -0.0043   1.0000   0.3142
  -6.250  -0.5631   0.08642   0.08151  -0.0362   1.0000   0.1559
  -6.000  -0.5624   0.08009   0.07507  -0.0397   1.0000   0.1547
  -5.750  -0.5573   0.07207   0.06686  -0.0462   1.0000   0.1538
  -5.500  -0.5313   0.05772   0.05151  -0.0635   1.0000   0.1560
  -5.250  -0.5248   0.05991   0.05412  -0.0570   1.0000   0.1647
  -5.000  -0.4940   0.05311   0.04667  -0.0652   1.0000   0.1750
  -4.750  -0.4634   0.04906   0.04201  -0.0703   1.0000   0.1889
  -4.500  -0.4369   0.04689   0.03947  -0.0727   1.0000   0.2041
  -4.250  -0.4127   0.04550   0.03785  -0.0739   1.0000   0.2200
  -4.000  -0.3904   0.04462   0.03684  -0.0744   1.0000   0.2367
  -3.750  -0.3690   0.04402   0.03615  -0.0745   1.0000   0.2542
  -3.500  -0.3483   0.04364   0.03573  -0.0744   1.0000   0.2726
  -3.250  -0.3278   0.04342   0.03546  -0.0742   1.0000   0.2925
  -3.000  -0.3049   0.04313   0.03502  -0.0747   1.0000   0.3157
  -2.750  -0.2831   0.04298   0.03476  -0.0748   1.0000   0.3386
  -2.500  -0.2644   0.04311   0.03491  -0.0741   1.0000   0.3603
  -2.250  -0.2406   0.04310   0.03470  -0.0748   1.0000   0.3877
  -2.000  -0.2233   0.04336   0.03503  -0.0737   1.0000   0.4094
  -1.750  -0.2044   0.04361   0.03526  -0.0731   1.0000   0.4336
  -1.500  -0.1838   0.04387   0.03544  -0.0731   1.0000   0.4597
  -1.250  -0.1661   0.04425   0.03583  -0.0722   1.0000   0.4833
  -1.000  -0.1451   0.04464   0.03612  -0.0723   1.0000   0.5107
  -0.750  -0.1283   0.04508   0.03660  -0.0712   1.0000   0.5342
  -0.500  -0.1077   0.04559   0.03702  -0.0713   1.0000   0.5621
  -0.250  -0.0915   0.04610   0.03757  -0.0701   1.0000   0.5861
   0.000  -0.0718   0.04671   0.03812  -0.0700   1.0000   0.6144
   0.250  -0.0553   0.04731   0.03873  -0.0690   1.0000   0.6401
   0.500  -0.0373   0.04798   0.03936  -0.0685   1.0000   0.6688
   0.750  -0.0194   0.04868   0.04003  -0.0679   1.0000   0.6998
   1.000  -0.0031   0.04936   0.04071  -0.0670   1.0000   0.7324
   1.250   0.0114   0.04997   0.04137  -0.0656   1.0000   0.7680
   1.500   0.0239   0.05046   0.04195  -0.0637   0.9998   0.8106
   1.750   0.0406   0.05108   0.04274  -0.0623   0.9948   0.8821
   2.000   0.0847   0.05242   0.04400  -0.0699   0.9833   1.0000
   2.250   0.1284   0.05447   0.04570  -0.0772   0.9749   1.0000
   2.500   0.1679   0.05701   0.04791  -0.0826   0.9685   1.0000
   2.750   0.2039   0.05931   0.04996  -0.0868   0.9584   1.0000
   3.000   0.2287   0.06083   0.05132  -0.0888   0.9478   1.0000
   3.250   0.2561   0.06294   0.05327  -0.0912   0.9396   1.0000
   3.500   0.2910   0.06560   0.05577  -0.0946   0.9292   1.0000
   3.750   0.3091   0.06682   0.05692  -0.0953   0.9174   1.0000
   4.000   0.3328   0.06890   0.05891  -0.0968   0.9084   1.0000
   4.250   0.3665   0.07165   0.06156  -0.0998   0.8976   1.0000
   4.500   0.3810   0.07283   0.06271  -0.0999   0.8853   1.0000
   4.750   0.4030   0.07503   0.06486  -0.1012   0.8761   1.0000
   5.000   0.4355   0.07782   0.06759  -0.1038   0.8649   1.0000
   5.250   0.4470   0.07903   0.06881  -0.1035   0.8523   1.0000
   5.500   0.4670   0.08128   0.07104  -0.1044   0.8426   1.0000
   5.750   0.4999   0.08429   0.07404  -0.1071   0.8315   1.0000
   6.000   0.5094   0.08555   0.07532  -0.1065   0.8184   1.0000
   6.250   0.5253   0.08770   0.07749  -0.1070   0.8076   1.0000
   6.500   0.5629   0.09138   0.08118  -0.1101   0.7976   1.0000
   6.750   0.5697   0.09252   0.08237  -0.1093   0.7837   1.0000
   7.000   0.5806   0.09451   0.08439  -0.1092   0.7718   1.0000
   7.250   0.6067   0.09768   0.08760  -0.1109   0.7623   1.0000
   7.500   0.6303   0.10021   0.09018  -0.1121   0.7487   1.0000
   7.750   0.6362   0.10187   0.09192  -0.1115   0.7353   1.0000
   8.000   0.6482   0.10427   0.09438  -0.1117   0.7234   1.0000
   8.250   0.6759   0.10772   0.09790  -0.1135   0.7131   1.0000
   8.500   0.6989   0.11041   0.10069  -0.1145   0.6986   1.0000
   8.750   0.7016   0.11215   0.10250  -0.1139   0.6849   1.0000
   9.000   0.7103   0.11460   0.10503  -0.1139   0.6721   1.0000
   9.250   0.7266   0.11761   0.10813  -0.1147   0.6608   1.0000
   9.500   0.7543   0.12105   0.11169  -0.1162   0.6477   1.0000
   9.750   0.7793   0.12412   0.11486  -0.1172   0.6322   1.0000
  10.000   0.7756   0.12579   0.11662  -0.1165   0.6184   1.0000
  10.250   0.7810   0.12837   0.11929  -0.1166   0.6052   1.0000
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