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

EPPLER 557 AIRFOIL (e557-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 557 AIRFOIL (e557-il)
Reynolds number: 100,000
Max Cl/Cd: 54.91 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e557-il-100000-n5.txt
Download as CSV file: xf-e557-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 557 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
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.5030   0.09574   0.09026  -0.0618   1.0000   0.0217
 -12.750  -0.5320   0.08661   0.08102  -0.0665   1.0000   0.0209
 -12.500  -0.5610   0.07876   0.07302  -0.0706   1.0000   0.0207
 -12.250  -0.5870   0.07224   0.06634  -0.0736   1.0000   0.0203
 -12.000  -0.6056   0.06750   0.06146  -0.0753   1.0000   0.0204
 -11.750  -0.6233   0.06315   0.05696  -0.0762   1.0000   0.0203
 -11.500  -0.6382   0.05961   0.05331  -0.0765   1.0000   0.0206
 -11.250  -0.6524   0.05654   0.05011  -0.0762   1.0000   0.0209
 -11.000  -0.6655   0.05379   0.04725  -0.0752   1.0000   0.0210
 -10.750  -0.6812   0.05131   0.04465  -0.0736   1.0000   0.0214
 -10.500  -0.6975   0.04928   0.04249  -0.0713   1.0000   0.0218
 -10.250  -0.7143   0.04758   0.04072  -0.0682   1.0000   0.0219
 -10.000  -0.7326   0.04590   0.03891  -0.0651   1.0000   0.0222
  -9.750  -0.7182   0.04350   0.03637  -0.0676   0.9958   0.0231
  -9.500  -0.6944   0.04134   0.03409  -0.0707   0.9902   0.0243
  -9.250  -0.6694   0.03919   0.03174  -0.0729   0.9849   0.0256
  -9.000  -0.6456   0.03733   0.02963  -0.0740   0.9790   0.0270
  -8.750  -0.6191   0.03570   0.02802  -0.0765   0.9734   0.0289
  -8.500  -0.5955   0.03422   0.02634  -0.0776   0.9662   0.0321
  -8.250  -0.5671   0.03269   0.02486  -0.0801   0.9611   0.0350
  -8.000  -0.5448   0.03136   0.02343  -0.0807   0.9530   0.0387
  -7.750  -0.5140   0.02996   0.02200  -0.0834   0.9480   0.0441
  -7.500  -0.4911   0.02867   0.02071  -0.0845   0.9397   0.0498
  -7.250  -0.4592   0.02729   0.01930  -0.0873   0.9346   0.0587
  -7.000  -0.4327   0.02602   0.01800  -0.0890   0.9271   0.0697
  -6.750  -0.4001   0.02465   0.01662  -0.0919   0.9214   0.0872
  -6.500  -0.3620   0.02299   0.01506  -0.0963   0.9179   0.1171
  -6.250  -0.3341   0.02134   0.01366  -0.0989   0.9095   0.1684
  -6.000  -0.2959   0.01984   0.01295  -0.1033   0.9053   0.3038
  -5.750  -0.2572   0.01984   0.01292  -0.1054   0.9014   0.3639
  -5.500  -0.2277   0.01992   0.01280  -0.1058   0.8929   0.3918
  -5.250  -0.1903   0.02012   0.01287  -0.1072   0.8886   0.4113
  -5.000  -0.1578   0.02025   0.01284  -0.1079   0.8819   0.4264
  -4.750  -0.1238   0.02032   0.01272  -0.1090   0.8754   0.4406
  -4.500  -0.0864   0.02066   0.01303  -0.1100   0.8714   0.4544
  -4.250  -0.0602   0.02099   0.01334  -0.1091   0.8627   0.4645
  -4.000  -0.0209   0.02071   0.01283  -0.1116   0.8574   0.4744
  -3.750   0.0147   0.02055   0.01260  -0.1128   0.8519   0.4780
  -3.500   0.0448   0.02038   0.01233  -0.1132   0.8435   0.4822
  -3.250   0.0859   0.01997   0.01172  -0.1160   0.8387   0.4877
  -3.000   0.1154   0.01973   0.01134  -0.1167   0.8295   0.4919
  -2.750   0.1521   0.01952   0.01105  -0.1182   0.8234   0.4951
  -2.500   0.1818   0.01938   0.01083  -0.1186   0.8146   0.4988
  -2.250   0.2186   0.01910   0.01042  -0.1205   0.8075   0.5031
  -2.000   0.2506   0.01889   0.01007  -0.1216   0.7987   0.5080
  -1.750   0.2843   0.01875   0.00989  -0.1226   0.7910   0.5110
  -1.500   0.3135   0.01867   0.00975  -0.1228   0.7817   0.5143
  -1.250   0.3482   0.01849   0.00949  -0.1242   0.7740   0.5183
  -1.000   0.3777   0.01840   0.00929  -0.1248   0.7642   0.5231
  -0.750   0.4118   0.01827   0.00910  -0.1259   0.7566   0.5267
  -0.500   0.4378   0.01827   0.00910  -0.1255   0.7462   0.5301
  -0.250   0.4705   0.01818   0.00895  -0.1264   0.7382   0.5341
   0.000   0.4986   0.01816   0.00887  -0.1266   0.7280   0.5384
   0.250   0.5288   0.01812   0.00878  -0.1271   0.7191   0.5427
   0.500   0.5571   0.01812   0.00879  -0.1271   0.7098   0.5463
   0.750   0.5847   0.01815   0.00881  -0.1271   0.7001   0.5507
   1.000   0.6158   0.01813   0.00872  -0.1278   0.6916   0.5557
   1.250   0.6421   0.01819   0.00879  -0.1275   0.6814   0.5597
   1.500   0.6715   0.01821   0.00880  -0.1277   0.6731   0.5637
   1.750   0.6970   0.01830   0.00892  -0.1273   0.6629   0.5686
   2.000   0.7259   0.01836   0.00895  -0.1276   0.6539   0.5744
   2.250   0.7523   0.01843   0.00907  -0.1273   0.6444   0.5785
   2.500   0.7784   0.01854   0.00921  -0.1270   0.6349   0.5832
   2.750   0.8073   0.01862   0.00928  -0.1271   0.6260   0.5889
   3.000   0.8321   0.01876   0.00948  -0.1266   0.6159   0.5941
   3.250   0.8597   0.01886   0.00960  -0.1265   0.6071   0.5996
   3.500   0.8849   0.01902   0.00981  -0.1261   0.5968   0.6061
   3.750   0.9107   0.01916   0.01001  -0.1257   0.5872   0.6116
   4.000   0.9367   0.01930   0.01019  -0.1253   0.5777   0.6176
   4.250   0.9615   0.01950   0.01044  -0.1249   0.5671   0.6250
   4.500   0.9870   0.01965   0.01064  -0.1243   0.5576   0.6311
   4.750   1.0111   0.01985   0.01091  -0.1237   0.5467   0.6390
   5.000   1.0346   0.02005   0.01121  -0.1229   0.5359   0.6460
   5.250   1.0598   0.02024   0.01141  -0.1224   0.5257   0.6543
   5.500   1.0820   0.02047   0.01174  -0.1214   0.5142   0.6624
   5.750   1.1046   0.02071   0.01208  -0.1204   0.5027   0.6719
   6.000   1.1273   0.02094   0.01237  -0.1195   0.4917   0.6812
   6.250   1.1493   0.02120   0.01270  -0.1184   0.4800   0.6920
   6.500   1.1694   0.02148   0.01310  -0.1171   0.4679   0.7028
   6.750   1.1896   0.02176   0.01349  -0.1157   0.4559   0.7147
   7.000   1.2096   0.02205   0.01385  -0.1143   0.4440   0.7281
   7.250   1.2278   0.02236   0.01426  -0.1126   0.4313   0.7432
   7.500   1.2442   0.02269   0.01474  -0.1106   0.4180   0.7601
   7.750   1.2593   0.02303   0.01520  -0.1084   0.4048   0.7799
   8.250   1.2814   0.02365   0.01605  -0.1024   0.3788   0.8397
   8.500   1.2844   0.02376   0.01631  -0.0978   0.3666   1.0000
   8.750   1.2989   0.02441   0.01697  -0.0960   0.3514   1.0000
   9.000   1.3118   0.02512   0.01769  -0.0940   0.3360   1.0000
   9.250   1.3231   0.02591   0.01847  -0.0919   0.3202   1.0000
   9.500   1.3329   0.02678   0.01935  -0.0896   0.3042   1.0000
   9.750   1.3415   0.02776   0.02033  -0.0873   0.2882   1.0000
  10.000   1.3485   0.02885   0.02141  -0.0850   0.2722   1.0000
  10.250   1.3545   0.03007   0.02262  -0.0827   0.2565   1.0000
  10.500   1.3593   0.03143   0.02397  -0.0804   0.2411   1.0000
  10.750   1.3633   0.03292   0.02546  -0.0783   0.2263   1.0000
  11.000   1.3667   0.03454   0.02710  -0.0763   0.2122   1.0000
  11.250   1.3695   0.03630   0.02889  -0.0744   0.1989   1.0000
  11.500   1.3717   0.03820   0.03081  -0.0727   0.1866   1.0000
  11.750   1.3730   0.04027   0.03289  -0.0712   0.1752   1.0000
  12.000   1.3726   0.04256   0.03517  -0.0697   0.1650   1.0000
  12.250   1.3740   0.04483   0.03752  -0.0686   0.1543   1.0000
  12.500   1.3742   0.04730   0.04004  -0.0676   0.1449   1.0000
  12.750   1.3719   0.05008   0.04282  -0.0667   0.1367   1.0000
  13.000   1.3728   0.05271   0.04557  -0.0661   0.1279   1.0000
  13.250   1.3705   0.05574   0.04863  -0.0656   0.1209   1.0000
  13.500   1.3698   0.05872   0.05173  -0.0654   0.1134   1.0000
  13.750   1.3669   0.06204   0.05509  -0.0653   0.1073   1.0000
  14.000   1.3658   0.06531   0.05847  -0.0653   0.1008   1.0000
  14.250   1.3614   0.06902   0.06219  -0.0656   0.0958   1.0000
  14.500   1.3608   0.07245   0.06578  -0.0660   0.0898   1.0000
  14.750   1.3561   0.07644   0.06981  -0.0666   0.0853   1.0000
  15.000   1.3542   0.08019   0.07368  -0.0673   0.0806   1.0000
  15.250   1.3509   0.08423   0.07783  -0.0682   0.0761   1.0000
  15.750   1.3442   0.09254   0.08635  -0.0704   0.0685   1.0000
  16.000   1.3402   0.09693   0.09085  -0.0718   0.0650   1.0000
  16.500   1.3335   0.10565   0.09976  -0.0748   0.0589   1.0000
  16.750   1.3301   0.11016   0.10440  -0.0766   0.0561   1.0000
<< Back to EPPLER 557 AIRFOIL (e557-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 557 AIRFOIL (e557-il)