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

EPPLER 580 AIRFOIL (e580-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 580 AIRFOIL (e580-il)
Reynolds number: 100,000
Max Cl/Cd: 54.39 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e580-il-100000-n5.txt
Download as CSV file: xf-e580-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 580 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  -0.4090   0.08185   0.07700  -0.0707   0.9912   0.0303
  -9.750  -0.4128   0.07528   0.07035  -0.0775   0.9854   0.0301
  -9.500  -0.4216   0.06923   0.06418  -0.0837   0.9773   0.0298
  -9.250  -0.4316   0.06406   0.05885  -0.0890   0.9677   0.0296
  -9.000  -0.4415   0.05960   0.05419  -0.0931   0.9563   0.0294
  -8.750  -0.4401   0.05477   0.04906  -0.0978   0.9461   0.0292
  -8.500  -0.4385   0.05083   0.04481  -0.0996   0.9350   0.0291
  -8.250  -0.4302   0.04720   0.04083  -0.1012   0.9256   0.0290
  -8.000  -0.4115   0.04346   0.03666  -0.1038   0.9196   0.0290
  -7.750  -0.3966   0.04057   0.03338  -0.1043   0.9108   0.0291
  -7.500  -0.3703   0.03768   0.03007  -0.1063   0.9059   0.0292
  -7.250  -0.3471   0.03536   0.02739  -0.1069   0.8992   0.0295
  -7.000  -0.3203   0.03334   0.02503  -0.1077   0.8933   0.0299
  -6.750  -0.2877   0.03142   0.02289  -0.1094   0.8899   0.0311
  -6.500  -0.2652   0.03025   0.02167  -0.1092   0.8821   0.0323
  -6.250  -0.2345   0.02892   0.02022  -0.1102   0.8774   0.0336
  -6.000  -0.2002   0.02752   0.01868  -0.1114   0.8744   0.0345
  -5.750  -0.1783   0.02657   0.01766  -0.1105   0.8660   0.0353
  -5.500  -0.1464   0.02549   0.01650  -0.1113   0.8616   0.0364
  -5.250  -0.1102   0.02444   0.01536  -0.1130   0.8586   0.0377
  -5.000  -0.0894   0.02366   0.01459  -0.1123   0.8492   0.0393
  -4.750  -0.0530   0.02280   0.01369  -0.1144   0.8451   0.0431
  -4.500  -0.0236   0.02212   0.01293  -0.1153   0.8381   0.0481
  -4.250   0.0103   0.02133   0.01207  -0.1171   0.8321   0.0552
  -4.000   0.0516   0.02016   0.01100  -0.1205   0.8283   0.0845
  -3.750   0.0862   0.01787   0.01007  -0.1253   0.8204   0.3481
  -3.500   0.1190   0.01833   0.01096  -0.1253   0.8147   0.5163
  -3.250   0.1434   0.01941   0.01204  -0.1230   0.8069   0.5653
  -3.000   0.1690   0.02029   0.01287  -0.1209   0.8000   0.5946
  -2.750   0.1966   0.02060   0.01306  -0.1199   0.7933   0.6072
  -2.500   0.2270   0.02047   0.01277  -0.1205   0.7858   0.6119
  -2.250   0.2622   0.02029   0.01242  -0.1218   0.7803   0.6156
  -2.000   0.2899   0.02019   0.01217  -0.1221   0.7716   0.6201
  -1.750   0.3307   0.01986   0.01161  -0.1251   0.7662   0.6253
  -1.500   0.3546   0.01986   0.01154  -0.1244   0.7573   0.6278
  -1.250   0.3890   0.01971   0.01125  -0.1257   0.7513   0.6308
  -1.000   0.4157   0.01967   0.01113  -0.1258   0.7427   0.6342
  -0.750   0.4509   0.01952   0.01082  -0.1275   0.7364   0.6384
  -0.500   0.4797   0.01947   0.01067  -0.1281   0.7283   0.6418
  -0.250   0.5103   0.01941   0.01053  -0.1287   0.7215   0.6442
   0.000   0.5375   0.01942   0.01049  -0.1287   0.7139   0.6471
   0.250   0.5678   0.01940   0.01039  -0.1294   0.7068   0.6505
   0.500   0.5981   0.01938   0.01029  -0.1303   0.6997   0.6540
   0.750   0.6279   0.01938   0.01022  -0.1310   0.6921   0.6574
   1.000   0.6566   0.01941   0.01020  -0.1312   0.6856   0.6599
   1.250   0.6823   0.01949   0.01028  -0.1310   0.6777   0.6629
   1.500   0.7139   0.01949   0.01021  -0.1318   0.6716   0.6660
   1.750   0.7392   0.01960   0.01032  -0.1318   0.6633   0.6695
   2.000   0.7719   0.01961   0.01025  -0.1329   0.6572   0.6731
   2.250   0.7941   0.01977   0.01047  -0.1320   0.6490   0.6758
   2.500   0.8235   0.01982   0.01049  -0.1324   0.6425   0.6789
   2.750   0.8481   0.01998   0.01069  -0.1321   0.6347   0.6822
   3.000   0.8778   0.02006   0.01073  -0.1327   0.6278   0.6859
   3.250   0.9033   0.02022   0.01093  -0.1325   0.6202   0.6894
   3.500   0.9296   0.02033   0.01108  -0.1323   0.6130   0.6925
   3.750   0.9550   0.02050   0.01129  -0.1320   0.6057   0.6961
   4.000   0.9815   0.02065   0.01146  -0.1320   0.5980   0.7000
   4.250   1.0087   0.02083   0.01165  -0.1322   0.5907   0.7040
   4.500   1.0325   0.02100   0.01191  -0.1316   0.5828   0.7072
   4.750   1.0575   0.02118   0.01214  -0.1313   0.5753   0.7111
   5.000   1.0821   0.02139   0.01240  -0.1309   0.5671   0.7157
   5.250   1.1074   0.02160   0.01268  -0.1307   0.5592   0.7202
   5.500   1.1307   0.02180   0.01296  -0.1300   0.5509   0.7239
   5.750   1.1534   0.02204   0.01328  -0.1293   0.5425   0.7282
   6.000   1.1787   0.02225   0.01353  -0.1290   0.5341   0.7334
   6.250   1.1987   0.02254   0.01395  -0.1278   0.5250   0.7381
   6.500   1.2241   0.02271   0.01414  -0.1274   0.5169   0.7433
   6.750   1.2420   0.02307   0.01464  -0.1260   0.5066   0.7494
   7.000   1.2627   0.02332   0.01498  -0.1249   0.4974   0.7547
   7.250   1.2830   0.02359   0.01534  -0.1237   0.4875   0.7606
   7.500   1.2999   0.02396   0.01583  -0.1220   0.4768   0.7673
   7.750   1.3174   0.02425   0.01621  -0.1204   0.4668   0.7741
   8.000   1.3344   0.02456   0.01659  -0.1186   0.4562   0.7823
   8.250   1.3460   0.02498   0.01717  -0.1161   0.4449   0.7907
   8.500   1.3600   0.02540   0.01770  -0.1139   0.4338   0.8002
   8.750   1.3743   0.02582   0.01819  -0.1119   0.4226   0.8111
   9.000   1.3853   0.02631   0.01879  -0.1093   0.4108   0.8232
   9.250   1.3943   0.02688   0.01950  -0.1065   0.3986   0.8389
   9.500   1.4019   0.02745   0.02021  -0.1036   0.3866   0.8610
   9.750   1.4058   0.02792   0.02081  -0.0999   0.3753   0.9104
  10.000   1.4145   0.02866   0.02159  -0.0976   0.3625   1.0000
  10.250   1.4242   0.02969   0.02264  -0.0958   0.3489   1.0000
  10.500   1.4318   0.03085   0.02385  -0.0939   0.3349   1.0000
  10.750   1.4382   0.03213   0.02517  -0.0919   0.3209   1.0000
  11.000   1.4433   0.03354   0.02664  -0.0899   0.3066   1.0000
  11.250   1.4472   0.03510   0.02823  -0.0879   0.2924   1.0000
  11.500   1.4497   0.03683   0.02999  -0.0860   0.2780   1.0000
  11.750   1.4513   0.03872   0.03191  -0.0842   0.2636   1.0000
  12.000   1.4517   0.04080   0.03402  -0.0825   0.2494   1.0000
  12.250   1.4511   0.04307   0.03632  -0.0810   0.2353   1.0000
  12.500   1.4498   0.04551   0.03878  -0.0796   0.2216   1.0000
  12.750   1.4478   0.04815   0.04145  -0.0784   0.2085   1.0000
  13.000   1.4451   0.05098   0.04430  -0.0774   0.1958   1.0000
  13.250   1.4416   0.05399   0.04736  -0.0766   0.1840   1.0000
  13.500   1.4376   0.05721   0.05061  -0.0760   0.1722   1.0000
  13.750   1.4338   0.06056   0.05402  -0.0756   0.1605   1.0000
  14.000   1.4291   0.06413   0.05764  -0.0754   0.1494   1.0000
  14.250   1.4231   0.06797   0.06152  -0.0755   0.1387   1.0000
  14.500   1.4158   0.07212   0.06566  -0.0757   0.1287   1.0000
  14.750   1.4090   0.07637   0.06996  -0.0762   0.1182   1.0000
  15.000   1.4021   0.08075   0.07441  -0.0769   0.1081   1.0000
  15.250   1.3941   0.08542   0.07911  -0.0778   0.0987   1.0000
  15.500   1.3846   0.09044   0.08413  -0.0790   0.0905   1.0000
  15.750   1.3782   0.09515   0.08893  -0.0803   0.0818   1.0000
<< Back to EPPLER 580 AIRFOIL (e580-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 580 AIRFOIL (e580-il)