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

EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 59 AIRFOIL (e59-il)
Reynolds number: 50,000
Max Cl/Cd: 46.21 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e59-il-50000-n5.txt
Download as CSV file: xf-e59-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 59 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
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.3922   0.10600   0.09959  -0.0185   1.0000   0.0774
  -6.500  -0.3931   0.10375   0.09740  -0.0179   1.0000   0.0792
  -6.250  -0.3936   0.10142   0.09514  -0.0180   1.0000   0.0813
  -6.000  -0.3933   0.09925   0.09304  -0.0197   1.0000   0.0839
  -5.750  -0.3880   0.09729   0.09115  -0.0261   1.0000   0.0857
  -5.500  -0.3786   0.09386   0.08778  -0.0302   1.0000   0.0866
  -5.250  -0.3790   0.09004   0.08400  -0.0228   1.0000   0.0900
  -5.000  -0.3706   0.08716   0.08114  -0.0241   1.0000   0.0958
  -4.750  -0.3471   0.08340   0.07737  -0.0352   1.0000   0.1011
  -4.500  -0.3448   0.08020   0.07421  -0.0307   1.0000   0.1044
  -4.250  -0.3286   0.07678   0.07078  -0.0335   1.0000   0.1093
  -3.750  -0.2347   0.06221   0.05570  -0.0584   1.0000   0.0559
  -3.500  -0.2011   0.05745   0.05078  -0.0643   1.0000   0.0540
  -3.250  -0.1394   0.05088   0.04376  -0.0776   1.0000   0.0558
  -3.000  -0.0760   0.04463   0.03691  -0.0896   1.0000   0.0561
  -2.750  -0.0065   0.03868   0.03003  -0.1015   1.0000   0.0579
  -2.500   0.0262   0.03673   0.02796  -0.1043   1.0000   0.0654
  -2.250   0.0824   0.03305   0.02330  -0.1112   1.0000   0.0693
  -2.000   0.1186   0.03147   0.02134  -0.1138   1.0000   0.0799
  -1.750   0.1581   0.02971   0.01886  -0.1163   1.0000   0.0853
  -1.500   0.1904   0.02857   0.01748  -0.1175   1.0000   0.0908
  -1.250   0.2204   0.02794   0.01653  -0.1182   1.0000   0.1020
  -0.750   0.2767   0.02718   0.01543  -0.1188   1.0000   0.1251
  -0.500   0.3052   0.02694   0.01505  -0.1192   1.0000   0.1407
  -0.250   0.3406   0.02675   0.01493  -0.1213   0.9973   0.1869
   0.000   0.3846   0.02585   0.01511  -0.1253   0.9951   0.4271
   0.250   0.4051   0.02496   0.01515  -0.1235   0.9871   1.0000
   0.500   0.4397   0.02550   0.01533  -0.1252   0.9806   1.0000
   0.750   0.4733   0.02602   0.01559  -0.1269   0.9739   1.0000
   1.000   0.5096   0.02659   0.01594  -0.1290   0.9680   1.0000
   1.250   0.5413   0.02705   0.01628  -0.1303   0.9600   1.0000
   1.500   0.5740   0.02753   0.01666  -0.1317   0.9525   1.0000
   1.750   0.6090   0.02802   0.01710  -0.1335   0.9453   1.0000
   2.000   0.6391   0.02845   0.01751  -0.1345   0.9361   1.0000
   2.250   0.6765   0.02890   0.01796  -0.1366   0.9294   1.0000
   2.500   0.7055   0.02928   0.01841  -0.1373   0.9188   1.0000
   2.750   0.7367   0.02964   0.01884  -0.1383   0.9085   1.0000
   3.000   0.7717   0.02982   0.01913  -0.1397   0.8967   1.0000
   3.250   0.8070   0.02982   0.01929  -0.1409   0.8822   1.0000
   3.500   0.8434   0.02961   0.01925  -0.1419   0.8658   1.0000
   3.750   0.8815   0.02922   0.01906  -0.1430   0.8493   1.0000
   4.000   0.9186   0.02882   0.01894  -0.1439   0.8336   1.0000
   4.250   0.9460   0.02875   0.01911  -0.1433   0.8151   1.0000
   4.500   0.9800   0.02816   0.01881  -0.1431   0.7948   1.0000
   4.750   1.0137   0.02707   0.01806  -0.1420   0.7650   1.0000
   5.000   1.0449   0.02610   0.01733  -0.1403   0.7227   1.0000
   5.250   1.0811   0.02502   0.01638  -0.1391   0.6546   1.0000
   5.500   1.1344   0.02455   0.01492  -0.1398   0.4744   1.0000
   5.750   1.1446   0.02613   0.01582  -0.1364   0.3415   1.0000
   6.000   1.1539   0.02813   0.01722  -0.1336   0.2445   1.0000
   6.250   1.1658   0.03041   0.01889  -0.1314   0.1454   1.0000
   6.500   1.1815   0.03250   0.02071  -0.1296   0.1132   1.0000
   6.750   1.1996   0.03433   0.02255  -0.1282   0.0969   1.0000
   7.000   1.2193   0.03619   0.02456  -0.1269   0.0853   1.0000
   7.250   1.2435   0.03813   0.02673  -0.1260   0.0779   1.0000
   7.500   1.2712   0.04032   0.02913  -0.1258   0.0680   1.0000
   7.750   1.3006   0.04298   0.03184  -0.1259   0.0597   1.0000
   8.000   1.3277   0.04547   0.03471  -0.1255   0.0520   1.0000
   8.250   1.3558   0.04900   0.03848  -0.1254   0.0467   1.0000
   8.500   1.3772   0.05232   0.04230  -0.1242   0.0423   1.0000
   8.750   1.3947   0.05604   0.04648  -0.1228   0.0398   1.0000
   9.000   1.4057   0.06041   0.05114  -0.1211   0.0375   1.0000
   9.250   1.4072   0.06422   0.05568  -0.1174   0.0359   1.0000
   9.500   1.4047   0.06836   0.06039  -0.1138   0.0348   1.0000
   9.750   1.3977   0.07265   0.06516  -0.1101   0.0344   1.0000
  10.000   1.3850   0.07671   0.06962  -0.1061   0.0342   1.0000
  10.250   1.3687   0.08082   0.07408  -0.1024   0.0341   1.0000
  10.500   1.3503   0.08519   0.07877  -0.0993   0.0341   1.0000
  10.750   1.3304   0.08990   0.08377  -0.0972   0.0342   1.0000
  11.000   1.3092   0.09507   0.08920  -0.0962   0.0343   1.0000
  11.250   1.2873   0.10080   0.09516  -0.0965   0.0345   1.0000
  11.500   1.2654   0.10720   0.10175  -0.0983   0.0348   1.0000
<< Back to EPPLER 59 AIRFOIL (e59-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 59 AIRFOIL (e59-il)