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

EPPLER 398 AIRFOIL (e398-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 398 AIRFOIL (e398-il)
Reynolds number: 50,000
Max Cl/Cd: 17.74 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e398-il-50000-n5.txt
Download as CSV file: xf-e398-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 398 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.3038   0.12364   0.11717  -0.0490   1.0000   0.0536
 -10.500  -0.3153   0.12150   0.11513  -0.0475   1.0000   0.0533
 -10.250  -0.3269   0.11967   0.11340  -0.0456   1.0000   0.0538
 -10.000  -0.3155   0.11473   0.10846  -0.0504   0.9941   0.0550
  -9.750  -0.3082   0.10939   0.10313  -0.0554   0.9871   0.0569
  -9.500  -0.3036   0.10367   0.09743  -0.0607   0.9800   0.0578
  -9.250  -0.3017   0.09750   0.09129  -0.0662   0.9722   0.0585
  -9.000  -0.3059   0.09099   0.08483  -0.0715   0.9627   0.0590
  -8.750  -0.3153   0.08360   0.07750  -0.0776   0.9528   0.0594
  -8.500  -0.3337   0.07499   0.06892  -0.0849   0.9414   0.0595
  -8.250  -0.3518   0.06508   0.05888  -0.0964   0.9284   0.0591
  -8.000  -0.3659   0.05642   0.04979  -0.1060   0.9155   0.0594
  -7.750  -0.3572   0.05363   0.04690  -0.1078   0.9058   0.0624
  -7.500  -0.3379   0.04922   0.04208  -0.1130   0.8995   0.0668
  -7.250  -0.3282   0.04479   0.03701  -0.1160   0.8899   0.0717
  -7.000  -0.2992   0.04282   0.03484  -0.1185   0.8848   0.0791
  -6.750  -0.2781   0.04064   0.03233  -0.1199   0.8777   0.0862
  -6.500  -0.2505   0.03841   0.02961  -0.1220   0.8716   0.0964
  -6.250  -0.2153   0.03687   0.02785  -0.1246   0.8678   0.1080
  -6.000  -0.1968   0.03607   0.02701  -0.1240   0.8596   0.1168
  -5.750  -0.1652   0.03486   0.02555  -0.1257   0.8548   0.1301
  -5.500  -0.1280   0.03372   0.02421  -0.1281   0.8515   0.1465
  -5.250  -0.1110   0.03324   0.02361  -0.1271   0.8428   0.1593
  -5.000  -0.0780   0.03249   0.02270  -0.1286   0.8384   0.1775
  -4.750  -0.0400   0.03170   0.02166  -0.1309   0.8353   0.2000
  -4.500  -0.0249   0.03163   0.02152  -0.1294   0.8265   0.2159
  -4.250   0.0078   0.03124   0.02111  -0.1305   0.8222   0.2371
  -4.000   0.0402   0.03088   0.02062  -0.1316   0.8181   0.2607
  -3.750   0.0579   0.03094   0.02060  -0.1305   0.8102   0.2803
  -3.500   0.0899   0.03071   0.02034  -0.1314   0.8061   0.3035
  -3.250   0.1199   0.03056   0.02003  -0.1320   0.8015   0.3284
  -3.000   0.1381   0.03076   0.02022  -0.1308   0.7942   0.3480
  -2.750   0.1698   0.03062   0.02002  -0.1316   0.7902   0.3728
  -2.500   0.1990   0.03060   0.01990  -0.1320   0.7858   0.3974
  -2.250   0.2160   0.03093   0.02023  -0.1306   0.7783   0.4175
  -2.000   0.2478   0.03087   0.02007  -0.1313   0.7744   0.4445
  -1.750   0.2779   0.03084   0.02002  -0.1316   0.7707   0.4692
  -1.500   0.2906   0.03140   0.02057  -0.1298   0.7625   0.4898
  -1.250   0.3209   0.03138   0.02054  -0.1301   0.7586   0.5166
  -0.750   0.3602   0.03211   0.02128  -0.1280   0.7467   0.5664
  -0.500   0.3896   0.03213   0.02130  -0.1281   0.7429   0.5967
  -0.250   0.4220   0.03201   0.02118  -0.1284   0.7401   0.6294
   0.000   0.4232   0.03302   0.02228  -0.1249   0.7306   0.6536
   0.250   0.4503   0.03304   0.02234  -0.1245   0.7269   0.6897
   0.500   0.4794   0.03286   0.02221  -0.1239   0.7242   0.7299
   0.750   0.4736   0.03402   0.02350  -0.1194   0.7142   0.7636
   1.000   0.4953   0.03388   0.02344  -0.1176   0.7106   0.8207
   1.500   0.5291   0.03507   0.02461  -0.1157   0.6972   1.0000
   1.750   0.5670   0.03524   0.02458  -0.1177   0.6942   1.0000
   2.000   0.5752   0.03677   0.02602  -0.1165   0.6853   1.0000
   2.250   0.6039   0.03729   0.02642  -0.1172   0.6807   1.0000
   2.500   0.6401   0.03739   0.02640  -0.1185   0.6779   1.0000
   2.750   0.6418   0.03931   0.02829  -0.1166   0.6678   1.0000
   3.000   0.6724   0.03967   0.02857  -0.1172   0.6639   1.0000
   3.250   0.6966   0.04038   0.02923  -0.1172   0.6588   1.0000
   3.500   0.7056   0.04195   0.03078  -0.1160   0.6501   1.0000
   3.750   0.7388   0.04211   0.03090  -0.1167   0.6469   1.0000
   4.250   0.7693   0.04458   0.03337  -0.1152   0.6325   1.0000
   4.500   0.7974   0.04500   0.03379  -0.1153   0.6283   1.0000
   4.750   0.8003   0.04703   0.03584  -0.1137   0.6182   1.0000
   5.000   0.8343   0.04704   0.03588  -0.1142   0.6150   1.0000
   5.250   0.8323   0.04944   0.03832  -0.1123   0.6038   1.0000
   5.500   0.8641   0.04955   0.03846  -0.1126   0.6000   1.0000
   6.000   0.8946   0.05203   0.04106  -0.1110   0.5849   1.0000
   6.500   0.9259   0.05443   0.04359  -0.1095   0.5695   1.0000
   7.000   0.9576   0.05677   0.04610  -0.1080   0.5539   1.0000
   7.500   0.9908   0.05891   0.04843  -0.1064   0.5381   1.0000
   7.750   0.9921   0.06127   0.05088  -0.1051   0.5264   1.0000
   8.000   1.0250   0.06085   0.05060  -0.1049   0.5220   1.0000
   8.500   1.0409   0.06440   0.05437  -0.1028   0.5010   1.0000
   8.750   1.0603   0.06504   0.05514  -0.1020   0.4927   1.0000
   9.250   1.0978   0.06629   0.05669  -0.1002   0.4759   1.0000
   9.500   1.0995   0.06869   0.05922  -0.0992   0.4630   1.0000
  10.000   1.1392   0.06933   0.06019  -0.0972   0.4456   1.0000
  10.250   1.1427   0.07153   0.06253  -0.0962   0.4324   1.0000
  10.750   1.1873   0.07099   0.06235  -0.0938   0.4145   1.0000
  11.000   1.1921   0.07296   0.06448  -0.0928   0.4009   1.0000
  11.500   1.2129   0.07540   0.06724  -0.0907   0.3755   1.0000
  11.750   1.2557   0.07191   0.06396  -0.0888   0.3678   1.0000
  12.000   1.2656   0.07306   0.06528  -0.0877   0.3533   1.0000
  12.500   1.2935   0.07415   0.06663  -0.0852   0.3228   1.0000
  12.750   1.3046   0.07515   0.06771  -0.0841   0.3059   1.0000
  13.000   1.3086   0.07726   0.06991  -0.0832   0.2881   1.0000
  13.250   1.3150   0.07904   0.07177  -0.0824   0.2699   1.0000
  13.500   1.3221   0.08071   0.07345  -0.0815   0.2513   1.0000
  13.750   1.3288   0.08246   0.07514  -0.0807   0.2329   1.0000
  14.000   1.3291   0.08536   0.07801  -0.0803   0.2156   1.0000
  14.250   1.3271   0.08874   0.08141  -0.0802   0.1989   1.0000
  14.500   1.3253   0.09218   0.08485  -0.0803   0.1834   1.0000
  14.750   1.3230   0.09577   0.08843  -0.0805   0.1687   1.0000
  15.000   1.3208   0.09942   0.09209  -0.0809   0.1552   1.0000
  15.250   1.3187   0.10313   0.09577  -0.0814   0.1428   1.0000
  15.500   1.3174   0.10671   0.09929  -0.0819   0.1314   1.0000
  15.750   1.3138   0.11100   0.10370  -0.0829   0.1207   1.0000
<< Back to EPPLER 398 AIRFOIL (e398-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 398 AIRFOIL (e398-il)