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EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 59 AIRFOIL (e59-il)
Reynolds number: 100,000
Max Cl/Cd: 69.92 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e59-il-100000.txt
Download as CSV file: xf-e59-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.500  -0.4147   0.09832   0.09381  -0.0143   1.0000   0.0629
  -6.250  -0.4147   0.09590   0.09144  -0.0140   1.0000   0.0646
  -6.000  -0.4136   0.09339   0.08895  -0.0146   1.0000   0.0666
  -5.750  -0.4099   0.09082   0.08642  -0.0170   1.0000   0.0690
  -5.500  -0.3844   0.08811   0.08370  -0.0339   1.0000   0.0715
  -5.250  -0.3779   0.08329   0.07890  -0.0343   1.0000   0.0724
  -5.000  -0.3802   0.08004   0.07571  -0.0279   1.0000   0.0739
  -4.750  -0.3730   0.07714   0.07283  -0.0265   1.0000   0.0765
  -4.500  -0.3566   0.07376   0.06943  -0.0298   1.0000   0.0810
  -4.250  -0.3168   0.06812   0.06368  -0.0433   1.0000   0.0878
  -4.000  -0.3111   0.06584   0.06144  -0.0399   1.0000   0.0926
  -3.750  -0.2671   0.06049   0.05593  -0.0516   1.0000   0.1027
  -3.500  -0.2200   0.05577   0.05099  -0.0622   1.0000   0.1157
  -3.250  -0.2088   0.05331   0.04859  -0.0599   1.0000   0.1206
  -3.000  -0.1477   0.04855   0.04345  -0.0732   1.0000   0.1441
  -2.750  -0.1230   0.04571   0.04061  -0.0746   1.0000   0.1600
  -2.500  -0.0874   0.04294   0.03771  -0.0788   1.0000   0.1882
  -2.250   0.0420   0.03136   0.02437  -0.1024   1.0000   0.1021
  -2.000   0.0894   0.02838   0.02074  -0.1066   1.0000   0.0978
  -1.750   0.1330   0.02590   0.01770  -0.1099   1.0000   0.0954
  -1.500   0.1695   0.02458   0.01590  -0.1117   1.0000   0.1023
  -1.250   0.2015   0.02358   0.01472  -0.1126   1.0000   0.1066
  -1.000   0.2315   0.02298   0.01391  -0.1132   1.0000   0.1158
  -0.750   0.2600   0.02251   0.01342  -0.1135   1.0000   0.1256
  -0.500   0.2890   0.02208   0.01305  -0.1139   1.0000   0.1378
  -0.250   0.3178   0.02184   0.01293  -0.1145   1.0000   0.1651
   0.000   0.3622   0.01965   0.01286  -0.1172   1.0000   0.7529
   0.250   0.3718   0.01969   0.01287  -0.1140   1.0000   1.0000
   0.500   0.3950   0.02020   0.01316  -0.1137   1.0000   1.0000
   0.750   0.4177   0.02075   0.01355  -0.1133   1.0000   1.0000
   1.000   0.4398   0.02133   0.01403  -0.1130   1.0000   1.0000
   1.250   0.4615   0.02197   0.01458  -0.1127   1.0000   1.0000
   1.500   0.4857   0.02268   0.01523  -0.1130   0.9984   1.0000
   1.750   0.5310   0.02351   0.01601  -0.1172   0.9875   1.0000
   2.000   0.5797   0.02417   0.01666  -0.1218   0.9746   1.0000
   2.250   0.6292   0.02457   0.01706  -0.1262   0.9609   1.0000
   2.500   0.6750   0.02480   0.01734  -0.1298   0.9479   1.0000
   2.750   0.7199   0.02487   0.01752  -0.1330   0.9349   1.0000
   3.000   0.7658   0.02473   0.01748  -0.1362   0.9215   1.0000
   3.250   0.8187   0.02388   0.01678  -0.1397   0.9035   1.0000
   3.500   0.8732   0.02246   0.01558  -0.1428   0.8830   1.0000
   3.750   0.9182   0.02156   0.01490  -0.1447   0.8682   1.0000
   4.000   0.9640   0.02000   0.01358  -0.1458   0.8469   1.0000
   4.250   1.0108   0.01775   0.01159  -0.1458   0.8121   1.0000
   4.500   1.0796   0.01544   0.00920  -0.1494   0.6947   1.0000
   4.750   1.1129   0.01718   0.00911  -0.1484   0.3785   1.0000
   5.000   1.1130   0.02073   0.01073  -0.1438   0.1352   1.0000
   5.250   1.1294   0.02235   0.01215  -0.1415   0.1081   1.0000
   5.500   1.1468   0.02377   0.01352  -0.1395   0.0959   1.0000
   5.750   1.1648   0.02545   0.01512  -0.1376   0.0892   1.0000
   6.000   1.1889   0.02706   0.01682  -0.1365   0.0837   1.0000
   6.250   1.2161   0.02923   0.01882  -0.1365   0.0759   1.0000
   6.500   1.2491   0.03152   0.02134  -0.1367   0.0722   1.0000
   6.750   1.2769   0.03346   0.02345  -0.1363   0.0662   1.0000
   7.000   1.3112   0.03741   0.02743  -0.1373   0.0626   1.0000
   7.250   1.3364   0.04073   0.03124  -0.1361   0.0613   1.0000
   7.500   1.3549   0.04311   0.03424  -0.1336   0.0586   1.0000
   7.750   1.3732   0.04706   0.03874  -0.1313   0.0586   1.0000
   8.000   1.3882   0.05197   0.04419  -0.1289   0.0601   1.0000
   8.250   1.3963   0.05710   0.05027  -0.1242   0.0668   1.0000
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