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

EPPLER 585 AIRFOIL (e585-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 585 AIRFOIL (e585-il)
Reynolds number: 500,000
Max Cl/Cd: 110.48 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e585-il-500000-n5.txt
Download as CSV file: xf-e585-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 585 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.500  -0.0462   0.10314   0.10036  -0.1010   0.8730   0.0060
 -12.250  -0.0422   0.09987   0.09704  -0.1022   0.8632   0.0059
 -12.000  -0.1573   0.11095   0.10839  -0.0901   0.9377   0.0060
 -11.500  -0.1169   0.10037   0.09771  -0.1035   0.9205   0.0056
 -11.250  -0.0986   0.09520   0.09245  -0.1100   0.9075   0.0053
 -11.000  -0.0865   0.09016   0.08732  -0.1151   0.8905   0.0049
 -10.750  -0.0829   0.08539   0.08246  -0.1182   0.8736   0.0049
 -10.500  -0.0847   0.08067   0.07767  -0.1205   0.8580   0.0051
 -10.250  -0.0884   0.07608   0.07303  -0.1224   0.8454   0.0053
 -10.000  -0.0942   0.07130   0.06821  -0.1245   0.8344   0.0053
  -9.750  -0.1022   0.06629   0.06317  -0.1268   0.8243   0.0050
  -9.500  -0.1264   0.05753   0.05443  -0.1327   0.8140   0.0049
  -9.250  -0.1778   0.04675   0.04347  -0.1378   0.8028   0.0046
  -8.750  -0.2311   0.03680   0.03306  -0.1330   0.7869   0.0047
  -8.500  -0.2567   0.02985   0.02556  -0.1293   0.7805   0.0048
  -8.250  -0.2555   0.02640   0.02170  -0.1271   0.7757   0.0049
  -8.000  -0.2476   0.02351   0.01838  -0.1253   0.7708   0.0050
  -7.750  -0.2333   0.02137   0.01587  -0.1239   0.7665   0.0051
  -7.500  -0.2147   0.01988   0.01407  -0.1230   0.7628   0.0052
  -7.250  -0.1934   0.01876   0.01274  -0.1224   0.7591   0.0055
  -7.000  -0.1712   0.01777   0.01153  -0.1218   0.7552   0.0057
  -6.750  -0.1479   0.01697   0.01054  -0.1213   0.7515   0.0060
  -6.500  -0.1238   0.01630   0.00969  -0.1209   0.7482   0.0062
  -6.250  -0.0999   0.01552   0.00878  -0.1204   0.7450   0.0063
  -6.000  -0.0766   0.01468   0.00784  -0.1199   0.7414   0.0064
  -5.750  -0.0526   0.01403   0.00711  -0.1195   0.7378   0.0065
  -5.500  -0.0283   0.01345   0.00645  -0.1191   0.7346   0.0066
  -5.000   0.0220   0.01254   0.00543  -0.1187   0.7285   0.0071
  -4.750   0.0477   0.01216   0.00500  -0.1185   0.7251   0.0074
  -4.500   0.0738   0.01183   0.00461  -0.1184   0.7219   0.0079
  -4.250   0.1003   0.01154   0.00425  -0.1183   0.7189   0.0084
  -4.000   0.1268   0.01124   0.00389  -0.1183   0.7161   0.0097
  -3.750   0.1536   0.01098   0.00361  -0.1182   0.7128   0.0118
  -3.500   0.1805   0.01072   0.00337  -0.1183   0.7096   0.0181
  -3.250   0.2073   0.01045   0.00316  -0.1184   0.7065   0.0367
  -3.000   0.2341   0.01017   0.00299  -0.1185   0.7035   0.0666
  -2.750   0.2610   0.00987   0.00285  -0.1187   0.7005   0.1160
  -2.500   0.2877   0.00942   0.00271  -0.1190   0.6973   0.2031
  -2.250   0.3145   0.00894   0.00257  -0.1195   0.6939   0.3131
  -2.000   0.3416   0.00837   0.00241  -0.1200   0.6905   0.4529
  -1.750   0.3680   0.00780   0.00242  -0.1203   0.6874   0.6386
  -1.500   0.3957   0.00781   0.00250  -0.1202   0.6840   0.6877
  -1.250   0.4236   0.00785   0.00255  -0.1203   0.6800   0.7087
  -1.000   0.4517   0.00791   0.00256  -0.1204   0.6763   0.7227
  -0.750   0.4798   0.00798   0.00259  -0.1205   0.6728   0.7346
  -0.500   0.5075   0.00809   0.00268  -0.1204   0.6695   0.7506
  -0.250   0.5344   0.00820   0.00281  -0.1202   0.6655   0.7645
   0.000   0.5619   0.00827   0.00286  -0.1201   0.6617   0.7715
   0.250   0.5898   0.00831   0.00286  -0.1203   0.6580   0.7741
   0.500   0.6179   0.00836   0.00287  -0.1205   0.6542   0.7764
   0.750   0.6459   0.00838   0.00288  -0.1207   0.6495   0.7786
   1.000   0.6738   0.00843   0.00288  -0.1208   0.6446   0.7809
   1.250   0.7018   0.00849   0.00289  -0.1211   0.6401   0.7831
   1.500   0.7293   0.00851   0.00293  -0.1212   0.6348   0.7847
   1.750   0.7564   0.00856   0.00296  -0.1212   0.6296   0.7863
   2.000   0.7836   0.00862   0.00301  -0.1212   0.6248   0.7882
   2.250   0.8107   0.00867   0.00307  -0.1212   0.6188   0.7902
   2.500   0.8374   0.00874   0.00311  -0.1212   0.6127   0.7922
   2.750   0.8643   0.00881   0.00318  -0.1212   0.6060   0.7942
   3.000   0.8908   0.00889   0.00325  -0.1211   0.5991   0.7962
   3.250   0.9174   0.00897   0.00332  -0.1211   0.5920   0.7984
   3.500   0.9430   0.00906   0.00341  -0.1209   0.5841   0.8003
   3.750   0.9686   0.00915   0.00351  -0.1206   0.5762   0.8021
   4.000   0.9933   0.00927   0.00362  -0.1202   0.5676   0.8040
   4.250   1.0184   0.00938   0.00374  -0.1199   0.5579   0.8060
   4.500   1.0427   0.00952   0.00387  -0.1194   0.5479   0.8081
   4.750   1.0661   0.00969   0.00402  -0.1188   0.5365   0.8104
   5.000   1.0893   0.00987   0.00418  -0.1181   0.5239   0.8129
   5.250   1.1114   0.01006   0.00436  -0.1172   0.5100   0.8152
   5.500   1.1319   0.01028   0.00457  -0.1160   0.4947   0.8173
   5.750   1.1512   0.01055   0.00480  -0.1146   0.4784   0.8196
   6.000   1.1696   0.01083   0.00505  -0.1131   0.4617   0.8222
   6.250   1.1858   0.01114   0.00532  -0.1111   0.4439   0.8250
   6.500   1.2002   0.01149   0.00562  -0.1088   0.4259   0.8280
   6.750   1.2138   0.01189   0.00597  -0.1064   0.4076   0.8310
   7.000   1.2261   0.01234   0.00639  -0.1038   0.3888   0.8342
   7.250   1.2384   0.01282   0.00683  -0.1013   0.3695   0.8377
   7.500   1.2499   0.01338   0.00733  -0.0988   0.3492   0.8414
   7.750   1.2599   0.01401   0.00789  -0.0961   0.3290   0.8450
   8.000   1.2706   0.01460   0.00846  -0.0936   0.3122   0.8486
   8.250   1.2810   0.01526   0.00909  -0.0911   0.2940   0.8529
   8.500   1.2898   0.01603   0.00982  -0.0885   0.2742   0.8577
   8.750   1.2985   0.01683   0.01058  -0.0860   0.2568   0.8628
   9.000   1.3064   0.01770   0.01142  -0.0835   0.2396   0.8688
   9.500   1.3217   0.01960   0.01327  -0.0788   0.2049   0.8826
   9.750   1.3290   0.02061   0.01427  -0.0765   0.1889   0.8915
  10.000   1.3350   0.02169   0.01534  -0.0741   0.1729   0.9039
  10.250   1.3402   0.02281   0.01646  -0.0716   0.1574   0.9263
  10.750   1.3559   0.02534   0.01894  -0.0685   0.1281   1.0000
  11.000   1.3639   0.02669   0.02026  -0.0670   0.1154   1.0000
  11.250   1.3716   0.02811   0.02165  -0.0656   0.1034   1.0000
  11.500   1.3789   0.02959   0.02311  -0.0643   0.0919   1.0000
  11.750   1.3853   0.03117   0.02466  -0.0629   0.0804   1.0000
  12.000   1.3911   0.03285   0.02631  -0.0616   0.0695   1.0000
  12.250   1.3968   0.03458   0.02802  -0.0604   0.0601   1.0000
  12.500   1.4020   0.03639   0.02981  -0.0592   0.0511   1.0000
  12.750   1.4060   0.03837   0.03176  -0.0580   0.0421   1.0000
  13.000   1.4105   0.04036   0.03374  -0.0570   0.0353   1.0000
  13.250   1.4133   0.04255   0.03593  -0.0560   0.0278   1.0000
  13.500   1.4180   0.04464   0.03803  -0.0552   0.0235   1.0000
  13.750   1.4207   0.04696   0.04036  -0.0544   0.0183   1.0000
  14.000   1.4268   0.04902   0.04248  -0.0539   0.0162   1.0000
  14.250   1.4295   0.05148   0.04496  -0.0533   0.0126   1.0000
  14.500   1.4336   0.05384   0.04737  -0.0529   0.0109   1.0000
  14.750   1.4381   0.05622   0.04982  -0.0526   0.0094   1.0000
  15.000   1.4411   0.05883   0.05249  -0.0524   0.0080   1.0000
  15.250   1.4453   0.06135   0.05511  -0.0522   0.0071   1.0000
  15.500   1.4470   0.06423   0.05806  -0.0522   0.0059   1.0000
  15.750   1.4500   0.06701   0.06093  -0.0523   0.0052   1.0000
  16.000   1.4509   0.07011   0.06411  -0.0524   0.0045   1.0000
  16.250   1.4526   0.07318   0.06727  -0.0527   0.0041   1.0000
  16.500   1.4533   0.07646   0.07065  -0.0531   0.0035   1.0000
  16.750   1.4526   0.07999   0.07427  -0.0536   0.0031   1.0000
  17.000   1.4516   0.08361   0.07800  -0.0543   0.0027   1.0000
  17.250   1.4509   0.08728   0.08177  -0.0551   0.0025   1.0000
  17.500   1.4486   0.09123   0.08584  -0.0560   0.0022   1.0000
  17.750   1.4440   0.09563   0.09033  -0.0572   0.0018   1.0000
  18.000   1.4417   0.09973   0.09456  -0.0585   0.0018   1.0000
  18.250   1.4382   0.10408   0.09904  -0.0599   0.0015   1.0000
  18.500   1.4339   0.10863   0.10371  -0.0615   0.0014   1.0000
  18.750   1.4286   0.11339   0.10859  -0.0633   0.0013   1.0000
  19.000   1.4234   0.11822   0.11354  -0.0653   0.0012   1.0000
  19.250   1.4163   0.12342   0.11885  -0.0675   0.0011   1.0000
<< Back to EPPLER 585 AIRFOIL (e585-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 585 AIRFOIL (e585-il)