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=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 398 AIRFOIL (e398-il)
Reynolds number: 100,000
Max Cl/Cd: 54.92 at α=8.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e398-il-100000-n5.txt
Download as CSV file: xf-e398-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3969   0.05391   0.04880  -0.1070   0.9453   0.0243
  -9.750  -0.4078   0.04577   0.04028  -0.1186   0.9302   0.0244
  -9.500  -0.3969   0.04198   0.03633  -0.1246   0.9184   0.0253
  -9.250  -0.3808   0.03902   0.03313  -0.1288   0.9092   0.0267
  -9.000  -0.3638   0.03551   0.02916  -0.1326   0.9008   0.0286
  -8.750  -0.3454   0.03304   0.02639  -0.1343   0.8920   0.0303
  -8.500  -0.3192   0.03126   0.02439  -0.1367   0.8854   0.0338
  -8.250  -0.2959   0.02956   0.02248  -0.1380   0.8778   0.0373
  -8.000  -0.2696   0.02797   0.02055  -0.1394   0.8708   0.0422
  -7.750  -0.2408   0.02675   0.01926  -0.1411   0.8651   0.0476
  -7.500  -0.2168   0.02566   0.01802  -0.1415   0.8570   0.0536
  -7.250  -0.1853   0.02453   0.01666  -0.1432   0.8520   0.0617
  -7.000  -0.1611   0.02378   0.01577  -0.1433   0.8439   0.0695
  -6.750  -0.1315   0.02294   0.01484  -0.1444   0.8382   0.0783
  -6.500  -0.1032   0.02219   0.01396  -0.1451   0.8321   0.0879
  -6.250  -0.0765   0.02156   0.01321  -0.1454   0.8250   0.0985
  -6.000  -0.0448   0.02090   0.01242  -0.1465   0.8202   0.1110
  -5.750  -0.0195   0.02048   0.01189  -0.1465   0.8130   0.1236
  -5.500   0.0097   0.02001   0.01133  -0.1471   0.8072   0.1384
  -5.250   0.0403   0.01956   0.01083  -0.1480   0.8023   0.1548
  -5.000   0.0655   0.01928   0.01051  -0.1478   0.7951   0.1706
  -4.750   0.0956   0.01894   0.01012  -0.1485   0.7901   0.1887
  -4.500   0.1240   0.01870   0.00983  -0.1488   0.7847   0.2076
  -4.250   0.1506   0.01853   0.00961  -0.1488   0.7783   0.2264
  -4.000   0.1810   0.01832   0.00929  -0.1494   0.7736   0.2474
  -3.750   0.2084   0.01819   0.00915  -0.1495   0.7682   0.2666
  -3.500   0.2351   0.01810   0.00904  -0.1494   0.7623   0.2858
  -3.250   0.2651   0.01797   0.00883  -0.1499   0.7579   0.3067
  -3.000   0.2926   0.01791   0.00872  -0.1499   0.7527   0.3268
  -2.750   0.3189   0.01789   0.00869  -0.1498   0.7469   0.3463
  -2.500   0.3481   0.01780   0.00857  -0.1500   0.7426   0.3663
  -2.250   0.3767   0.01776   0.00850  -0.1502   0.7382   0.3866
  -2.000   0.4017   0.01780   0.00856  -0.1498   0.7323   0.4064
  -1.750   0.4300   0.01778   0.00849  -0.1499   0.7278   0.4274
  -1.500   0.4601   0.01771   0.00841  -0.1502   0.7243   0.4486
  -1.250   0.4833   0.01783   0.00858  -0.1496   0.7183   0.4686
  -1.000   0.5103   0.01786   0.00862  -0.1494   0.7135   0.4908
  -0.750   0.5396   0.01784   0.00859  -0.1496   0.7098   0.5139
  -0.500   0.5648   0.01794   0.00873  -0.1492   0.7050   0.5369
  -0.250   0.5897   0.01805   0.00889  -0.1487   0.6998   0.5613
   0.000   0.6172   0.01807   0.00894  -0.1486   0.6957   0.5869
   0.250   0.6466   0.01806   0.00894  -0.1487   0.6924   0.6143
   0.500   0.6678   0.01829   0.00927  -0.1476   0.6865   0.6413
   0.750   0.6928   0.01836   0.00942  -0.1470   0.6819   0.6702
   1.000   0.7201   0.01837   0.00946  -0.1466   0.6783   0.7017
   1.250   0.7423   0.01851   0.00969  -0.1455   0.6737   0.7345
   1.500   0.7624   0.01866   0.00995  -0.1440   0.6684   0.7702
   1.750   0.7844   0.01865   0.01003  -0.1425   0.6644   0.8114
   2.000   0.8074   0.01854   0.00997  -0.1410   0.6612   0.8643
   2.250   0.8272   0.01869   0.01025  -0.1397   0.6551   1.0000
   2.500   0.8556   0.01893   0.01044  -0.1401   0.6505   1.0000
   2.750   0.8871   0.01908   0.01051  -0.1410   0.6469   1.0000
   3.000   0.9123   0.01943   0.01086  -0.1408   0.6419   1.0000
   3.250   0.9366   0.01978   0.01121  -0.1405   0.6364   1.0000
   3.500   0.9659   0.01997   0.01136  -0.1410   0.6324   1.0000
   3.750   0.9951   0.02017   0.01155  -0.1413   0.6283   1.0000
   4.000   1.0148   0.02067   0.01211  -0.1403   0.6220   1.0000
   4.250   1.0427   0.02087   0.01232  -0.1404   0.6174   1.0000
   4.500   1.0751   0.02095   0.01237  -0.1412   0.6139   1.0000
   4.750   1.0905   0.02157   0.01311  -0.1395   0.6066   1.0000
   5.000   1.1177   0.02178   0.01334  -0.1395   0.6017   1.0000
   5.250   1.1473   0.02191   0.01349  -0.1398   0.5972   1.0000
   5.500   1.1633   0.02248   0.01418  -0.1382   0.5898   1.0000
   5.750   1.1923   0.02258   0.01430  -0.1383   0.5847   1.0000
   6.000   1.2124   0.02300   0.01482  -0.1372   0.5780   1.0000
   6.250   1.2347   0.02328   0.01518  -0.1364   0.5713   1.0000
   6.500   1.2627   0.02337   0.01531  -0.1363   0.5654   1.0000
   6.750   1.2774   0.02388   0.01596  -0.1344   0.5569   1.0000
   7.000   1.3069   0.02386   0.01596  -0.1345   0.5506   1.0000
   7.250   1.3181   0.02444   0.01669  -0.1320   0.5413   1.0000
   7.500   1.3424   0.02455   0.01688  -0.1313   0.5337   1.0000
   7.750   1.3565   0.02497   0.01742  -0.1292   0.5242   1.0000
   8.000   1.3698   0.02537   0.01792  -0.1268   0.5146   1.0000
   8.250   1.3932   0.02537   0.01795  -0.1259   0.5051   1.0000
   8.500   1.3962   0.02610   0.01882  -0.1222   0.4936   1.0000
   8.750   1.4055   0.02668   0.01951  -0.1194   0.4821   1.0000
   9.000   1.4173   0.02717   0.02007  -0.1171   0.4701   1.0000
   9.250   1.4285   0.02772   0.02068  -0.1148   0.4571   1.0000
   9.500   1.4366   0.02845   0.02148  -0.1122   0.4429   1.0000
   9.750   1.4422   0.02939   0.02248  -0.1095   0.4277   1.0000
  10.000   1.4460   0.03053   0.02369  -0.1068   0.4112   1.0000
  10.250   1.4490   0.03183   0.02502  -0.1042   0.3935   1.0000
  10.500   1.4517   0.03324   0.02645  -0.1017   0.3744   1.0000
  10.750   1.4540   0.03477   0.02794  -0.0994   0.3542   1.0000
  11.000   1.4538   0.03662   0.02977  -0.0971   0.3329   1.0000
  11.250   1.4524   0.03868   0.03179  -0.0949   0.3113   1.0000
  11.500   1.4496   0.04096   0.03402  -0.0929   0.2898   1.0000
  11.750   1.4460   0.04347   0.03648  -0.0910   0.2684   1.0000
  12.000   1.4413   0.04620   0.03913  -0.0894   0.2480   1.0000
  12.250   1.4369   0.04905   0.04195  -0.0879   0.2286   1.0000
  12.500   1.4323   0.05207   0.04493  -0.0867   0.2098   1.0000
  12.750   1.4276   0.05522   0.04805  -0.0857   0.1922   1.0000
  13.000   1.4229   0.05851   0.05131  -0.0849   0.1759   1.0000
  13.250   1.4184   0.06190   0.05468  -0.0843   0.1605   1.0000
  13.500   1.4142   0.06539   0.05815  -0.0838   0.1461   1.0000
  13.750   1.4104   0.06894   0.06171  -0.0835   0.1327   1.0000
  14.000   1.4069   0.07256   0.06536  -0.0834   0.1203   1.0000
  14.250   1.4037   0.07623   0.06907  -0.0834   0.1090   1.0000
  14.500   1.4006   0.07999   0.07286  -0.0835   0.0989   1.0000
  14.750   1.3964   0.08396   0.07685  -0.0838   0.0901   1.0000
  15.000   1.3928   0.08797   0.08090  -0.0842   0.0818   1.0000
  15.250   1.3904   0.09187   0.08488  -0.0847   0.0744   1.0000
  15.500   1.3851   0.09626   0.08927  -0.0855   0.0685   1.0000
  15.750   1.3836   0.10014   0.09329  -0.0862   0.0622   1.0000
<< Back to EPPLER 398 AIRFOIL (e398-il)

Polar data table (+)

Polar graphs


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