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

EPPLER 377 (MODIFIED) AIRFOIL (e377m-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 377 (MODIFIED) AIRFOIL (e377m-il)
Reynolds number: 500,000
Max Cl/Cd: 115.61 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e377m-il-500000-n5.txt
Download as CSV file: xf-e377m-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 377 (MODIFIED) 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.2129   0.11600   0.11367  -0.0118   0.8459   0.0051
 -10.250  -0.2064   0.11321   0.11077  -0.0125   0.8177   0.0051
 -10.000  -0.1998   0.11042   0.10788  -0.0134   0.7932   0.0052
  -9.750  -0.1926   0.10759   0.10497  -0.0145   0.7731   0.0052
  -9.500  -0.1852   0.10474   0.10205  -0.0155   0.7552   0.0052
  -9.250  -0.1778   0.10088   0.09811  -0.0160   0.7382   0.0054
  -9.000  -0.1694   0.09831   0.09548  -0.0169   0.7226   0.0058
  -8.750  -0.1615   0.09558   0.09269  -0.0178   0.7083   0.0060
  -8.250  -0.1449   0.09008   0.08710  -0.0197   0.6836   0.0070
  -8.000  -0.1366   0.08730   0.08429  -0.0206   0.6733   0.0072
  -7.750  -0.1285   0.08458   0.08153  -0.0214   0.6632   0.0075
  -7.500  -0.1202   0.08191   0.07883  -0.0223   0.6530   0.0077
  -7.250  -0.1125   0.07937   0.07627  -0.0230   0.6440   0.0078
  -7.000  -0.1052   0.07689   0.07377  -0.0237   0.6361   0.0079
  -6.750  -0.0966   0.07430   0.07116  -0.0247   0.6284   0.0080
  -6.500  -0.0856   0.07159   0.06842  -0.0263   0.6216   0.0080
  -6.250  -0.0730   0.06875   0.06556  -0.0282   0.6144   0.0081
  -6.000  -0.0594   0.06586   0.06262  -0.0303   0.6078   0.0081
  -5.750  -0.0442   0.06288   0.05963  -0.0326   0.6014   0.0081
  -5.500  -0.0287   0.05994   0.05666  -0.0349   0.5952   0.0081
  -5.250  -0.0112   0.05687   0.05356  -0.0375   0.5899   0.0082
  -5.000   0.0073   0.05376   0.05042  -0.0403   0.5839   0.0082
  -4.750   0.0272   0.05061   0.04723  -0.0432   0.5786   0.0082
  -4.500   0.0488   0.04733   0.04392  -0.0464   0.5736   0.0082
  -4.250   0.0712   0.04408   0.04063  -0.0496   0.5686   0.0082
  -4.000   0.0974   0.04083   0.03732  -0.0535   0.5642   0.0083
  -3.750   0.1185   0.03727   0.03373  -0.0561   0.5595   0.0087
  -3.500   0.1408   0.03539   0.03181  -0.0581   0.5539   0.0103
  -3.250   0.1767   0.03318   0.02951  -0.0630   0.5494   0.0126
  -3.000   0.2076   0.03037   0.02664  -0.0669   0.5452   0.0126
  -2.750   0.2392   0.02759   0.02379  -0.0707   0.5406   0.0127
  -2.500   0.2710   0.02493   0.02103  -0.0744   0.5364   0.0127
  -2.250   0.3052   0.02236   0.01838  -0.0782   0.5322   0.0128
  -1.750   0.3869   0.03669   0.03228  -0.0924   0.5311   0.0160
  -1.500   0.4253   0.03506   0.03054  -0.0956   0.5265   0.0190
  -1.250   0.4623   0.03336   0.02870  -0.0984   0.5222   0.0193
  -1.000   0.4915   0.03079   0.02608  -0.1005   0.5179   0.0175
  -0.750   0.5274   0.02865   0.02381  -0.1030   0.5134   0.0159
  -0.500   0.5620   0.02688   0.02192  -0.1050   0.5095   0.0162
  -0.250   0.5969   0.02531   0.02022  -0.1067   0.5055   0.0183
   0.000   0.6298   0.02371   0.01850  -0.1082   0.5010   0.0175
   0.250   0.6634   0.02207   0.01672  -0.1096   0.4968   0.0166
   0.500   0.6976   0.02030   0.01476  -0.1108   0.4932   0.0160
   0.750   0.7340   0.01793   0.01216  -0.1122   0.4893   0.0160
   1.000   0.7748   0.01312   0.00659  -0.1141   0.4856   0.0176
   1.250   0.8033   0.01228   0.00551  -0.1141   0.4813   0.0186
   1.500   0.8311   0.01184   0.00494  -0.1140   0.4774   0.0196
   1.750   0.8587   0.01172   0.00477  -0.1138   0.4731   0.0222
   2.000   0.8861   0.01146   0.00444  -0.1136   0.4686   0.0249
   2.250   0.9134   0.01094   0.00384  -0.1134   0.4646   0.0390
   2.500   0.9406   0.01090   0.00387  -0.1132   0.4603   0.0581
   2.750   0.9675   0.01100   0.00397  -0.1130   0.4557   0.0669
   3.250   1.0212   0.01135   0.00434  -0.1125   0.4470   0.0777
   3.500   1.0480   0.01133   0.00430  -0.1123   0.4420   0.0811
   4.000   1.1014   0.01121   0.00414  -0.1118   0.4313   0.0881
   4.250   1.1279   0.01129   0.00420  -0.1116   0.4249   0.0898
   4.500   1.1544   0.01129   0.00421  -0.1113   0.4183   0.0924
   4.750   1.1806   0.01117   0.00404  -0.1110   0.4114   0.0968
   5.000   1.2068   0.01116   0.00406  -0.1108   0.4048   0.0983
   5.250   1.2331   0.01125   0.00417  -0.1105   0.3976   0.0993
   5.500   1.2592   0.01134   0.00429  -0.1103   0.3904   0.1005
   5.750   1.2852   0.01147   0.00442  -0.1100   0.3819   0.1021
   6.000   1.3111   0.01158   0.00459  -0.1098   0.3729   0.1045
   6.250   1.3367   0.01169   0.00470  -0.1095   0.3632   0.1067
   6.500   1.3620   0.01183   0.00483  -0.1091   0.3521   0.1084
   6.750   1.3873   0.01200   0.00501  -0.1088   0.3393   0.1096
   7.000   1.4122   0.01224   0.00525  -0.1085   0.3236   0.1109
   7.250   1.4369   0.01253   0.00557  -0.1081   0.3063   0.1122
   7.500   1.4608   0.01292   0.00592  -0.1077   0.2847   0.1137
   7.750   1.4816   0.01375   0.00651  -0.1070   0.2377   0.1154
   8.000   1.4979   0.01521   0.00752  -0.1060   0.1687   0.1176
   8.250   1.5146   0.01652   0.00850  -0.1050   0.1172   0.1198
   8.500   1.5329   0.01752   0.00931  -0.1041   0.0853   0.1212
   8.750   1.5516   0.01838   0.01009  -0.1031   0.0646   0.1226
   9.000   1.5671   0.01957   0.01116  -0.1019   0.0375   0.1240
   9.250   1.5809   0.02084   0.01231  -0.1004   0.0170   0.1254
   9.500   1.5924   0.02223   0.01365  -0.0985   0.0045   0.1269
   9.750   1.6083   0.02306   0.01458  -0.0972   0.0034   0.1289
  10.000   1.6224   0.02396   0.01560  -0.0956   0.0029   0.1308
  10.250   1.6341   0.02495   0.01670  -0.0937   0.0026   0.1324
  10.500   1.6406   0.02602   0.01790  -0.0911   0.0024   0.1342
  10.750   1.6425   0.02730   0.01932  -0.0881   0.0022   0.1358
  11.000   1.6417   0.02898   0.02115  -0.0854   0.0021   0.1373
  11.250   1.6417   0.03088   0.02322  -0.0834   0.0020   0.1389
  11.500   1.6430   0.03291   0.02537  -0.0820   0.0019   0.1406
  11.750   1.6433   0.03524   0.02783  -0.0809   0.0019   0.1422
  12.000   1.6433   0.03781   0.03052  -0.0801   0.0018   0.1434
  12.250   1.6414   0.04076   0.03360  -0.0796   0.0017   0.1444
  12.500   1.6390   0.04391   0.03688  -0.0793   0.0017   0.1457
  12.750   1.6349   0.04739   0.04050  -0.0792   0.0016   0.1467
  13.000   1.6299   0.05108   0.04433  -0.0792   0.0016   0.1482
  13.250   1.6251   0.05485   0.04823  -0.0794   0.0016   0.1497
  13.500   1.6187   0.05896   0.05247  -0.0798   0.0015   0.1509
  13.750   1.6114   0.06327   0.05691  -0.0804   0.0015   0.1521
  14.000   1.6044   0.06766   0.06143  -0.0811   0.0015   0.1535
  14.250   1.5972   0.07226   0.06616  -0.0820   0.0015   0.1549
  14.500   1.5892   0.07719   0.07122  -0.0832   0.0014   0.1563
  14.750   1.5817   0.08217   0.07634  -0.0845   0.0014   0.1581
  15.000   1.5746   0.08721   0.08152  -0.0859   0.0014   0.1600
<< Back to EPPLER 377 (MODIFIED) AIRFOIL (e377m-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 377 (MODIFIED) AIRFOIL (e377m-il)