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EPPLER 59 AIRFOIL (e59-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 59 AIRFOIL (e59-il)
Reynolds number: 200,000
Max Cl/Cd: 103.63 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e59-il-200000.txt
Download as CSV file: xf-e59-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -5.750  -0.4154   0.08668   0.08362  -0.0238   1.0000   0.0335
  -5.500  -0.4183   0.08423   0.08120  -0.0175   1.0000   0.0348
  -5.250  -0.4119   0.08160   0.07857  -0.0174   1.0000   0.0360
  -5.000  -0.4014   0.07847   0.07544  -0.0196   1.0000   0.0373
  -4.750  -0.3863   0.07488   0.07185  -0.0233   1.0000   0.0388
  -4.500  -0.3570   0.07033   0.06725  -0.0312   0.9991   0.0413
  -4.250  -0.2492   0.06103   0.05752  -0.0625   0.9952   0.0445
  -4.000  -0.2343   0.05517   0.05176  -0.0641   0.9921   0.0465
  -3.750  -0.2015   0.05206   0.04860  -0.0682   0.9886   0.0494
  -3.500  -0.1147   0.04323   0.03917  -0.0885   0.9882   0.0590
  -3.250  -0.0865   0.04150   0.03753  -0.0903   0.9855   0.0638
  -3.000  -0.0214   0.03599   0.03150  -0.1017   0.9852   0.0746
  -2.750   0.0317   0.03273   0.02781  -0.1089   0.9844   0.0880
  -2.500   0.0985   0.02514   0.01908  -0.1166   0.9861   0.0551
  -2.250   0.1450   0.02250   0.01582  -0.1205   0.9857   0.0545
  -2.000   0.1869   0.02110   0.01400  -0.1234   0.9848   0.0603
  -1.750   0.2228   0.02021   0.01297  -0.1252   0.9818   0.0646
  -1.500   0.2581   0.01994   0.01251  -0.1267   0.9777   0.0714
  -1.250   0.2968   0.01899   0.01155  -0.1290   0.9757   0.0772
  -1.000   0.3357   0.01878   0.01130  -0.1313   0.9733   0.0865
  -0.750   0.3724   0.01835   0.01095  -0.1333   0.9698   0.0988
  -0.500   0.4078   0.01798   0.01072  -0.1351   0.9653   0.1293
  -0.250   0.4483   0.01693   0.01099  -0.1385   0.9638   0.5068
   0.000   0.4739   0.01608   0.01109  -0.1374   0.9605   1.0000
   0.250   0.5053   0.01621   0.01109  -0.1382   0.9545   1.0000
   0.500   0.5445   0.01625   0.01102  -0.1404   0.9492   1.0000
   0.750   0.5837   0.01623   0.01092  -0.1426   0.9434   1.0000
   1.000   0.6204   0.01616   0.01080  -0.1443   0.9367   1.0000
   1.250   0.6647   0.01604   0.01067  -0.1474   0.9334   1.0000
   1.500   0.6943   0.01594   0.01056  -0.1476   0.9244   1.0000
   1.750   0.7410   0.01536   0.01000  -0.1506   0.9170   1.0000
   2.000   0.7934   0.01431   0.00900  -0.1544   0.9074   1.0000
   2.250   0.8284   0.01379   0.00854  -0.1551   0.8967   1.0000
   2.500   0.8672   0.01310   0.00792  -0.1564   0.8868   1.0000
   2.750   0.9069   0.01212   0.00701  -0.1574   0.8728   1.0000
   3.000   0.9458   0.01114   0.00611  -0.1582   0.8472   1.0000
   3.250   0.9967   0.01033   0.00532  -0.1617   0.8060   1.0000
   3.500   1.0549   0.01018   0.00462  -0.1666   0.6752   1.0000
   3.750   1.0650   0.01157   0.00507  -0.1624   0.5250   1.0000
   4.000   1.0745   0.01292   0.00571  -0.1588   0.3798   1.0000
   4.250   1.0815   0.01490   0.00667  -0.1552   0.2118   1.0000
   4.500   1.0937   0.01693   0.00781  -0.1526   0.0809   1.0000
   4.750   1.1153   0.01776   0.00864  -0.1513   0.0696   1.0000
   5.000   1.1362   0.01864   0.00956  -0.1500   0.0640   1.0000
   5.250   1.1564   0.01961   0.01060  -0.1484   0.0605   1.0000
   5.500   1.1779   0.02042   0.01148  -0.1471   0.0561   1.0000
   5.750   1.1980   0.02155   0.01262  -0.1457   0.0527   1.0000
   6.000   1.2184   0.02331   0.01438  -0.1443   0.0495   1.0000
   6.250   1.2420   0.02442   0.01563  -0.1433   0.0459   1.0000
   6.500   1.2672   0.02601   0.01729  -0.1426   0.0430   1.0000
   6.750   1.2984   0.02973   0.02104  -0.1434   0.0386   1.0000
   7.000   1.3228   0.03108   0.02269  -0.1423   0.0368   1.0000
   7.250   1.3492   0.03370   0.02564  -0.1414   0.0355   1.0000
   7.500   1.3719   0.03636   0.02868  -0.1400   0.0338   1.0000
   7.750   1.3916   0.03844   0.03097  -0.1387   0.0312   1.0000
   8.000   1.4089   0.04231   0.03533  -0.1363   0.0311   1.0000
   8.250   1.4174   0.04926   0.04315  -0.1319   0.0352   1.0000
  13.500   0.9688   0.18211   0.17983  -0.1344   0.0495   1.0000
  13.750   0.9582   0.18836   0.18607  -0.1397   0.0476   1.0000
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