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

EPPLER 657 AIRFOIL (e657-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 657 AIRFOIL (e657-il)
Reynolds number: 500,000
Max Cl/Cd: 117.79 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e657-il-500000-n5.txt
Download as CSV file: xf-e657-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 657 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
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.2270   0.07235   0.06867  -0.1230   0.7687   0.0060
 -12.000  -0.2964   0.05378   0.04988  -0.1335   0.7666   0.0059
 -11.750  -0.3266   0.04453   0.04042  -0.1408   0.7639   0.0058
 -11.500  -0.3440   0.03984   0.03556  -0.1436   0.7612   0.0058
 -11.250  -0.3638   0.03595   0.03144  -0.1442   0.7588   0.0058
 -11.000  -0.3721   0.03356   0.02892  -0.1434   0.7567   0.0058
 -10.750  -0.3856   0.03098   0.02613  -0.1414   0.7544   0.0058
 -10.500  -0.3930   0.02911   0.02409  -0.1388   0.7522   0.0058
 -10.250  -0.3986   0.02761   0.02240  -0.1356   0.7501   0.0058
 -10.000  -0.3952   0.02600   0.02056  -0.1334   0.7482   0.0058
  -9.750  -0.3867   0.02457   0.01891  -0.1317   0.7465   0.0058
  -9.500  -0.3747   0.02327   0.01739  -0.1303   0.7448   0.0058
  -9.250  -0.3599   0.02210   0.01602  -0.1291   0.7431   0.0058
  -9.000  -0.3431   0.02103   0.01478  -0.1281   0.7416   0.0059
  -8.750  -0.3244   0.02009   0.01370  -0.1273   0.7401   0.0059
  -8.500  -0.3051   0.01914   0.01262  -0.1265   0.7387   0.0060
  -8.250  -0.2853   0.01822   0.01159  -0.1257   0.7372   0.0060
  -8.000  -0.2649   0.01733   0.01061  -0.1250   0.7358   0.0062
  -7.750  -0.2425   0.01669   0.00988  -0.1246   0.7344   0.0062
  -7.500  -0.2195   0.01609   0.00921  -0.1241   0.7330   0.0064
  -7.250  -0.1957   0.01556   0.00861  -0.1238   0.7315   0.0066
  -7.000  -0.1715   0.01507   0.00805  -0.1235   0.7300   0.0068
  -6.750  -0.1468   0.01461   0.00751  -0.1233   0.7286   0.0071
  -6.500  -0.1216   0.01420   0.00701  -0.1231   0.7274   0.0074
  -6.250  -0.0960   0.01382   0.00655  -0.1230   0.7262   0.0078
  -6.000  -0.0703   0.01344   0.00613  -0.1229   0.7251   0.0082
  -5.750  -0.0445   0.01307   0.00575  -0.1228   0.7238   0.0089
  -5.500  -0.0182   0.01277   0.00541  -0.1228   0.7224   0.0099
  -5.250   0.0083   0.01247   0.00509  -0.1228   0.7209   0.0111
  -5.000   0.0351   0.01221   0.00480  -0.1228   0.7194   0.0130
  -4.750   0.0619   0.01193   0.00454  -0.1229   0.7181   0.0171
  -4.500   0.0888   0.01167   0.00431  -0.1230   0.7168   0.0256
  -4.250   0.1160   0.01143   0.00411  -0.1231   0.7154   0.0366
  -4.000   0.1432   0.01120   0.00392  -0.1233   0.7141   0.0512
  -3.750   0.1706   0.01096   0.00375  -0.1235   0.7127   0.0716
  -3.500   0.1980   0.01069   0.00360  -0.1238   0.7115   0.1041
  -3.250   0.2256   0.01044   0.00348  -0.1241   0.7103   0.1439
  -3.000   0.2529   0.01021   0.00338  -0.1244   0.7089   0.1850
  -2.750   0.2803   0.00997   0.00329  -0.1247   0.7074   0.2284
  -2.500   0.3079   0.00974   0.00322  -0.1250   0.7059   0.2788
  -2.250   0.3354   0.00945   0.00314  -0.1254   0.7043   0.3464
  -2.000   0.3631   0.00914   0.00308  -0.1258   0.7026   0.4256
  -1.750   0.3910   0.00882   0.00303  -0.1262   0.7010   0.5102
  -1.500   0.4191   0.00859   0.00301  -0.1266   0.6994   0.5852
  -1.250   0.4472   0.00844   0.00302  -0.1268   0.6978   0.6436
  -1.000   0.4754   0.00838   0.00306  -0.1269   0.6961   0.6871
  -0.750   0.5042   0.00839   0.00308  -0.1272   0.6945   0.7167
  -0.500   0.5319   0.00840   0.00313  -0.1272   0.6925   0.7365
  -0.250   0.5590   0.00842   0.00322  -0.1271   0.6903   0.7525
   0.000   0.5863   0.00847   0.00330  -0.1270   0.6878   0.7665
   0.250   0.6138   0.00853   0.00338  -0.1270   0.6853   0.7790
   0.500   0.6418   0.00859   0.00343  -0.1271   0.6829   0.7902
   0.750   0.6702   0.00864   0.00346  -0.1273   0.6805   0.8003
   1.000   0.6996   0.00872   0.00347  -0.1277   0.6782   0.8099
   1.250   0.7263   0.00877   0.00357  -0.1276   0.6753   0.8165
   1.500   0.7542   0.00883   0.00363  -0.1278   0.6720   0.8240
   1.750   0.7807   0.00886   0.00368  -0.1276   0.6687   0.8290
   2.000   0.8084   0.00889   0.00370  -0.1276   0.6654   0.8341
   2.250   0.8372   0.00892   0.00369  -0.1280   0.6624   0.8384
   2.500   0.8629   0.00895   0.00378  -0.1277   0.6582   0.8415
   2.750   0.8893   0.00898   0.00384  -0.1276   0.6538   0.8450
   3.000   0.9164   0.00900   0.00387  -0.1276   0.6496   0.8487
   3.250   0.9436   0.00904   0.00391  -0.1277   0.6451   0.8523
   3.500   0.9694   0.00907   0.00398  -0.1274   0.6389   0.8553
   3.750   0.9949   0.00909   0.00401  -0.1271   0.6329   0.8581
   4.000   1.0202   0.00914   0.00410  -0.1268   0.6258   0.8610
   4.250   1.0452   0.00920   0.00414  -0.1264   0.6180   0.8641
   4.500   1.0703   0.00927   0.00424  -0.1261   0.6090   0.8672
   4.750   1.0943   0.00936   0.00432  -0.1255   0.5985   0.8701
   5.000   1.1156   0.00948   0.00441  -0.1244   0.5856   0.8728
   5.250   1.1355   0.00964   0.00454  -0.1230   0.5702   0.8760
   5.500   1.1537   0.00986   0.00473  -0.1214   0.5526   0.8796
   5.750   1.1694   0.01016   0.00495  -0.1193   0.5323   0.8833
   6.000   1.1797   0.01051   0.00522  -0.1161   0.5111   0.8867
   6.250   1.1860   0.01091   0.00556  -0.1122   0.4915   0.8901
   6.500   1.1933   0.01139   0.00598  -0.1087   0.4713   0.8937
   6.750   1.2002   0.01194   0.00646  -0.1052   0.4516   0.8979
   7.000   1.2055   0.01259   0.00704  -0.1017   0.4319   0.9019
   7.250   1.2099   0.01328   0.00769  -0.0981   0.4138   0.9058
   7.500   1.2149   0.01406   0.00843  -0.0948   0.3954   0.9101
   7.750   1.2213   0.01488   0.00921  -0.0920   0.3778   0.9145
   8.000   1.2267   0.01578   0.01007  -0.0891   0.3597   0.9191
   8.250   1.2325   0.01671   0.01097  -0.0863   0.3425   0.9248
   8.750   1.2442   0.01873   0.01291  -0.0812   0.3074   0.9430
   9.000   1.2565   0.01973   0.01390  -0.0801   0.2911   0.9662
   9.500   1.2746   0.02199   0.01605  -0.0769   0.2581   1.0000
   9.750   1.2841   0.02315   0.01716  -0.0754   0.2424   1.0000
  10.000   1.2939   0.02432   0.01827  -0.0740   0.2278   1.0000
  10.250   1.3036   0.02550   0.01942  -0.0726   0.2136   1.0000
  10.500   1.3132   0.02673   0.02060  -0.0713   0.1996   1.0000
  10.750   1.3229   0.02796   0.02179  -0.0700   0.1865   1.0000
  11.000   1.3321   0.02926   0.02305  -0.0687   0.1732   1.0000
  11.250   1.3414   0.03058   0.02432  -0.0675   0.1604   1.0000
  11.500   1.3498   0.03199   0.02569  -0.0662   0.1475   1.0000
  11.750   1.3594   0.03335   0.02702  -0.0651   0.1361   1.0000
  12.000   1.3681   0.03480   0.02844  -0.0640   0.1244   1.0000
  12.250   1.3770   0.03627   0.02988  -0.0630   0.1136   1.0000
  12.500   1.3861   0.03775   0.03135  -0.0620   0.1045   1.0000
  12.750   1.3942   0.03935   0.03293  -0.0611   0.0961   1.0000
  13.000   1.4040   0.04083   0.03443  -0.0603   0.0881   1.0000
  13.250   1.4123   0.04248   0.03608  -0.0594   0.0812   1.0000
  13.500   1.4195   0.04426   0.03785  -0.0586   0.0737   1.0000
  13.750   1.4285   0.04591   0.03952  -0.0579   0.0676   1.0000
  14.000   1.4351   0.04781   0.04143  -0.0572   0.0617   1.0000
  14.250   1.4426   0.04966   0.04330  -0.0565   0.0561   1.0000
  14.500   1.4491   0.05166   0.04533  -0.0560   0.0510   1.0000
  14.750   1.4545   0.05381   0.04748  -0.0554   0.0460   1.0000
  15.000   1.4607   0.05593   0.04964  -0.0549   0.0415   1.0000
  15.250   1.4657   0.05819   0.05193  -0.0545   0.0375   1.0000
  15.500   1.4699   0.06061   0.05438  -0.0542   0.0337   1.0000
  15.750   1.4753   0.06294   0.05677  -0.0539   0.0308   1.0000
  16.250   1.4824   0.06817   0.06210  -0.0536   0.0252   1.0000
  16.500   1.4843   0.07105   0.06503  -0.0536   0.0229   1.0000
  16.750   1.4871   0.07390   0.06795  -0.0537   0.0208   1.0000
  17.000   1.4885   0.07695   0.07106  -0.0539   0.0190   1.0000
  17.250   1.4906   0.08000   0.07420  -0.0542   0.0175   1.0000
  17.750   1.4914   0.08667   0.08102  -0.0552   0.0146   1.0000
  18.000   1.4896   0.09039   0.08480  -0.0559   0.0129   1.0000
  18.250   1.4895   0.09396   0.08847  -0.0567   0.0120   1.0000
  18.500   1.4872   0.09788   0.09247  -0.0577   0.0109   1.0000
  18.750   1.4857   0.10174   0.09642  -0.0587   0.0101   1.0000
  19.000   1.4837   0.10570   0.10048  -0.0599   0.0092   1.0000
  19.250   1.4802   0.10993   0.10480  -0.0613   0.0085   1.0000
<< Back to EPPLER 657 AIRFOIL (e657-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 657 AIRFOIL (e657-il)