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

EPPLER 560 AIRFOIL (e560-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 560 AIRFOIL (e560-il)
Reynolds number: 50,000
Max Cl/Cd: 31.07 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e560-il-50000-n5.txt
Download as CSV file: xf-e560-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 560 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.3403   0.11387   0.10648  -0.0499   1.0000   0.0735
 -11.000  -0.3536   0.10828   0.10097  -0.0515   1.0000   0.0755
 -10.750  -0.3805   0.10010   0.09292  -0.0544   1.0000   0.0773
 -10.500  -0.3665   0.10080   0.09367  -0.0521   1.0000   0.0803
 -10.250  -0.3721   0.09765   0.09060  -0.0520   1.0000   0.0831
 -10.000  -0.3947   0.09166   0.08474  -0.0533   1.0000   0.0857
  -9.750  -0.4465   0.08125   0.07452  -0.0566   1.0000   0.0873
  -9.500  -0.5019   0.07239   0.06579  -0.0585   1.0000   0.0870
  -9.250  -0.5507   0.06651   0.05999  -0.0580   1.0000   0.0861
  -9.000  -0.5966   0.06063   0.05412  -0.0587   1.0000   0.0852
  -8.750  -0.6270   0.05420   0.04751  -0.0628   1.0000   0.0859
  -8.500  -0.6279   0.04938   0.04245  -0.0677   0.9969   0.0905
  -8.250  -0.6005   0.04535   0.03811  -0.0757   0.9877   0.1016
  -8.000  -0.5687   0.04315   0.03578  -0.0812   0.9790   0.1149
  -7.750  -0.5337   0.04238   0.03503  -0.0853   0.9704   0.1302
  -7.500  -0.4995   0.04199   0.03463  -0.0883   0.9612   0.1457
  -7.250  -0.4675   0.04099   0.03346  -0.0913   0.9515   0.1600
  -7.000  -0.4341   0.03974   0.03197  -0.0945   0.9424   0.1738
  -6.750  -0.3970   0.03856   0.03053  -0.0981   0.9343   0.1879
  -6.500  -0.3677   0.03901   0.03105  -0.0978   0.9238   0.1971
  -6.250  -0.3332   0.03821   0.03006  -0.1002   0.9150   0.2090
  -6.000  -0.2981   0.03715   0.02875  -0.1029   0.9060   0.2219
  -5.750  -0.2650   0.03604   0.02731  -0.1055   0.8962   0.2352
  -5.500  -0.2286   0.03611   0.02742  -0.1064   0.8885   0.2442
  -5.250  -0.1988   0.03538   0.02649  -0.1077   0.8777   0.2553
  -5.000  -0.1563   0.03435   0.02516  -0.1112   0.8715   0.2688
  -4.750  -0.1300   0.03428   0.02508  -0.1108   0.8599   0.2770
  -4.500  -0.0903   0.03349   0.02408  -0.1133   0.8531   0.2886
  -4.250  -0.0572   0.03271   0.02302  -0.1151   0.8427   0.3012
  -4.000  -0.0232   0.03248   0.02279  -0.1158   0.8346   0.3095
  -3.750   0.0121   0.03184   0.02195  -0.1175   0.8254   0.3209
  -3.500   0.0469   0.03135   0.02130  -0.1189   0.8164   0.3317
  -3.250   0.0829   0.03093   0.02079  -0.1202   0.8079   0.3413
  -3.000   0.1171   0.03041   0.02005  -0.1218   0.7981   0.3531
  -2.750   0.1543   0.03001   0.01961  -0.1231   0.7902   0.3624
  -2.500   0.1846   0.02969   0.01914  -0.1237   0.7794   0.3727
  -2.250   0.2249   0.02916   0.01844  -0.1258   0.7722   0.3843
  -2.000   0.2510   0.02903   0.01828  -0.1255   0.7606   0.3928
  -1.750   0.2922   0.02851   0.01756  -0.1278   0.7535   0.4054
  -1.500   0.3170   0.02848   0.01752  -0.1272   0.7419   0.4137
  -1.250   0.3498   0.02824   0.01716  -0.1281   0.7325   0.4249
  -1.000   0.3830   0.02802   0.01684  -0.1289   0.7232   0.4361
  -0.750   0.4099   0.02800   0.01679  -0.1287   0.7128   0.4456
  -0.500   0.4465   0.02773   0.01638  -0.1301   0.7047   0.4582
  -0.250   0.4704   0.02783   0.01647  -0.1294   0.6936   0.4683
   0.000   0.5078   0.02757   0.01612  -0.1307   0.6865   0.4806
   0.250   0.5303   0.02778   0.01629  -0.1301   0.6752   0.4923
   0.500   0.5620   0.02772   0.01617  -0.1305   0.6670   0.5047
   0.750   0.5875   0.02784   0.01630  -0.1301   0.6572   0.5166
   1.000   0.6163   0.02791   0.01634  -0.1302   0.6488   0.5302
   1.250   0.6437   0.02803   0.01644  -0.1302   0.6398   0.5444
   1.500   0.6707   0.02818   0.01659  -0.1300   0.6313   0.5597
   1.750   0.6974   0.02831   0.01674  -0.1297   0.6229   0.5752
   2.000   0.7233   0.02850   0.01697  -0.1293   0.6149   0.5919
   2.250   0.7483   0.02870   0.01722  -0.1288   0.6065   0.6106
   2.500   0.7745   0.02885   0.01743  -0.1284   0.5991   0.6306
   2.750   0.7951   0.02917   0.01788  -0.1272   0.5907   0.6522
   3.000   0.8258   0.02913   0.01788  -0.1272   0.5848   0.6808
   3.250   0.8381   0.02969   0.01864  -0.1248   0.5757   0.7088
   3.500   0.8640   0.02960   0.01867  -0.1239   0.5699   0.7505
   3.750   0.8717   0.03006   0.01936  -0.1204   0.5619   0.8044
   4.000   0.8917   0.02998   0.01942  -0.1186   0.5552   1.0000
   4.250   0.9206   0.03058   0.01991  -0.1195   0.5483   1.0000
   4.500   0.9430   0.03136   0.02063  -0.1193   0.5406   1.0000
   4.750   0.9802   0.03155   0.02067  -0.1207   0.5355   1.0000
   5.000   0.9894   0.03283   0.02201  -0.1187   0.5268   1.0000
   5.250   1.0194   0.03324   0.02237  -0.1192   0.5210   1.0000
   5.500   1.0379   0.03416   0.02330  -0.1182   0.5143   1.0000
   5.750   1.0548   0.03513   0.02430  -0.1171   0.5072   1.0000
   6.000   1.0905   0.03530   0.02440  -0.1181   0.5024   1.0000
   6.250   1.0901   0.03705   0.02628  -0.1149   0.4941   1.0000
   6.500   1.1151   0.03766   0.02689  -0.1147   0.4885   1.0000
   6.750   1.1395   0.03833   0.02757  -0.1144   0.4831   1.0000
   7.000   1.1347   0.04028   0.02966  -0.1108   0.4750   1.0000
   7.250   1.1687   0.04049   0.02986  -0.1115   0.4703   1.0000
   7.500   1.1559   0.04289   0.03238  -0.1072   0.4626   1.0000
   7.750   1.1684   0.04410   0.03366  -0.1057   0.4563   1.0000
   8.000   1.2131   0.04368   0.03324  -0.1073   0.4524   1.0000
   8.250   1.1602   0.04878   0.03853  -0.1004   0.4418   1.0000
   8.500   1.1962   0.04856   0.03834  -0.1007   0.4376   1.0000
   8.750   1.1407   0.05512   0.04505  -0.0960   0.4259   1.0000
   9.000   1.1713   0.05506   0.04504  -0.0957   0.4217   1.0000
   9.250   1.2237   0.05332   0.04335  -0.0964   0.4192   1.0000
   9.500   1.1488   0.06265   0.05280  -0.0929   0.4048   1.0000
   9.750   1.1951   0.06081   0.05104  -0.0926   0.4028   1.0000
  11.500   1.1081   0.09463   0.08547  -0.0926   0.3368   1.0000
  11.750   1.0882   0.10116   0.09208  -0.0940   0.3253   1.0000
  12.000   1.1192   0.10006   0.09110  -0.0926   0.3225   1.0000
  12.250   1.0909   0.10799   0.09909  -0.0948   0.3096   1.0000
  12.500   1.1209   0.10696   0.09817  -0.0934   0.3068   1.0000
  13.000   1.1183   0.11460   0.10599  -0.0948   0.2902   1.0000
  13.250   1.0989   0.12151   0.11297  -0.0971   0.2787   1.0000
  13.500   1.1292   0.12007   0.11166  -0.0955   0.2756   1.0000
  13.750   1.1052   0.12795   0.11960  -0.0986   0.2634   1.0000
  14.000   1.1302   0.12734   0.11911  -0.0973   0.2597   1.0000
<< Back to EPPLER 560 AIRFOIL (e560-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 560 AIRFOIL (e560-il)