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 AIRFOIL (e377-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 377 AIRFOIL (e377-il)
Reynolds number: 500,000
Max Cl/Cd: 130.13 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e377-il-500000-n5.txt
Download as CSV file: xf-e377-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 377 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.1687   0.09776   0.09453  -0.0250   0.6865   0.0057
  -7.500  -0.1605   0.09549   0.09225  -0.0260   0.6767   0.0058
  -7.250  -0.1530   0.09236   0.08908  -0.0263   0.6668   0.0059
  -7.000  -0.1438   0.09013   0.08683  -0.0269   0.6572   0.0063
  -6.750  -0.1322   0.08786   0.08453  -0.0284   0.6476   0.0064
  -6.500  -0.1191   0.08569   0.08233  -0.0300   0.6404   0.0073
  -6.250  -0.1038   0.08337   0.07998  -0.0321   0.6316   0.0085
  -6.000  -0.0876   0.08100   0.07758  -0.0346   0.6245   0.0087
  -5.750  -0.0703   0.07854   0.07510  -0.0372   0.6178   0.0088
  -5.500  -0.0516   0.07602   0.07254  -0.0401   0.6109   0.0089
  -5.250  -0.0318   0.07343   0.06992  -0.0432   0.6041   0.0089
  -5.000  -0.0105   0.07079   0.06722  -0.0465   0.5977   0.0089
  -4.750   0.0122   0.06806   0.06447  -0.0499   0.5917   0.0090
  -4.500   0.0362   0.06530   0.06166  -0.0534   0.5860   0.0090
  -4.250   0.0615   0.06247   0.05879  -0.0571   0.5806   0.0090
  -4.000   0.0890   0.05951   0.05578  -0.0612   0.5747   0.0090
  -3.750   0.1179   0.05653   0.05273  -0.0653   0.5696   0.0090
  -3.500   0.1499   0.05351   0.04967  -0.0699   0.5647   0.0091
  -3.250   0.1761   0.05014   0.04626  -0.0733   0.5597   0.0096
  -3.000   0.2009   0.04861   0.04468  -0.0752   0.5547   0.0117
  -2.750   0.2451   0.04652   0.04249  -0.0813   0.5494   0.0140
  -2.500   0.2816   0.04390   0.03977  -0.0858   0.5444   0.0141
  -2.250   0.3180   0.04123   0.03702  -0.0900   0.5402   0.0141
  -2.000   0.3557   0.03852   0.03422  -0.0942   0.5356   0.0142
  -1.500   0.4226   0.03337   0.02889  -0.1007   0.5265   0.0160
  -1.250   0.4580   0.03146   0.02689  -0.1036   0.5221   0.0167
  -0.750   0.5398   0.02656   0.02167  -0.1107   0.5135   0.0138
  -0.500   0.5778   0.02430   0.01927  -0.1135   0.5094   0.0138
   0.000   0.6544   0.01957   0.01415  -0.1184   0.5011   0.0153
   0.250   0.6842   0.01885   0.01334  -0.1191   0.4970   0.0170
   0.500   0.7343   0.01149   0.00484  -0.1230   0.4938   0.0154
   0.750   0.7628   0.01084   0.00393  -0.1230   0.4894   0.0154
   1.000   0.7907   0.01042   0.00333  -0.1228   0.4854   0.0156
   1.250   0.8185   0.01011   0.00292  -0.1226   0.4813   0.0160
   1.500   0.8461   0.00991   0.00264  -0.1224   0.4766   0.0170
   1.750   0.8734   0.00979   0.00245  -0.1221   0.4724   0.0189
   2.000   0.9008   0.00973   0.00233  -0.1219   0.4685   0.0215
   2.250   0.9280   0.00964   0.00230  -0.1216   0.4639   0.0384
   2.500   0.9551   0.00967   0.00234  -0.1214   0.4594   0.0536
   2.750   0.9821   0.00972   0.00238  -0.1211   0.4553   0.0607
   3.000   1.0092   0.00976   0.00243  -0.1209   0.4507   0.0679
   3.250   1.0362   0.00981   0.00251  -0.1206   0.4458   0.0797
   3.500   1.0631   0.00988   0.00258  -0.1204   0.4413   0.0873
   3.750   1.0901   0.00993   0.00266  -0.1202   0.4357   0.0962
   4.000   1.1168   0.01002   0.00277  -0.1199   0.4297   0.1065
   4.250   1.1436   0.01009   0.00288  -0.1197   0.4234   0.1175
   4.500   1.1702   0.01020   0.00300  -0.1194   0.4164   0.1302
   4.750   1.1968   0.01030   0.00314  -0.1192   0.4098   0.1455
   5.000   1.2232   0.01042   0.00332  -0.1189   0.4026   0.1617
   5.250   1.2497   0.01053   0.00348  -0.1187   0.3956   0.1797
   5.500   1.2759   0.01068   0.00366  -0.1184   0.3874   0.1996
   5.750   1.3022   0.01081   0.00386  -0.1182   0.3786   0.2239
   6.000   1.3282   0.01098   0.00407  -0.1179   0.3692   0.2511
   6.250   1.3539   0.01116   0.00433  -0.1176   0.3582   0.2942
   6.500   1.3820   0.01062   0.00459  -0.1180   0.3456   1.0000
   6.750   1.4073   0.01090   0.00484  -0.1177   0.3317   1.0000
   7.000   1.4322   0.01123   0.00513  -0.1173   0.3150   1.0000
   7.250   1.4560   0.01172   0.00551  -0.1168   0.2872   1.0000
   7.500   1.4767   0.01269   0.00620  -0.1162   0.2329   1.0000
   7.750   1.4971   0.01368   0.00693  -0.1155   0.1887   1.0000
   8.000   1.5172   0.01467   0.00771  -0.1148   0.1509   1.0000
   8.250   1.5348   0.01593   0.00867  -0.1138   0.1045   1.0000
   8.500   1.5505   0.01737   0.00979  -0.1127   0.0611   1.0000
   8.750   1.5680   0.01847   0.01075  -0.1116   0.0375   1.0000
   9.000   1.5835   0.01972   0.01186  -0.1103   0.0164   1.0000
   9.250   1.5970   0.02111   0.01322  -0.1087   0.0042   1.0000
   9.500   1.6141   0.02199   0.01420  -0.1075   0.0033   1.0000
   9.750   1.6304   0.02287   0.01519  -0.1061   0.0030   1.0000
  10.000   1.6453   0.02380   0.01623  -0.1047   0.0029   1.0000
  10.250   1.6579   0.02483   0.01738  -0.1030   0.0027   1.0000
  10.500   1.6672   0.02594   0.01861  -0.1008   0.0025   1.0000
  10.750   1.6705   0.02720   0.01998  -0.0980   0.0023   1.0000
  11.000   1.6719   0.02873   0.02163  -0.0956   0.0022   1.0000
  11.250   1.6723   0.03063   0.02365  -0.0936   0.0020   1.0000
  11.500   1.6716   0.03291   0.02609  -0.0922   0.0019   1.0000
  11.750   1.6696   0.03561   0.02893  -0.0912   0.0018   1.0000
  12.000   1.6656   0.03876   0.03223  -0.0906   0.0018   1.0000
  12.250   1.6597   0.04234   0.03595  -0.0902   0.0017   1.0000
  12.500   1.6557   0.04580   0.03954  -0.0900   0.0017   1.0000
  12.750   1.6442   0.05035   0.04425  -0.0901   0.0016   1.0000
  13.000   1.6384   0.05425   0.04828  -0.0902   0.0016   1.0000
  13.250   1.6342   0.05801   0.05216  -0.0905   0.0016   1.0000
  13.500   1.6282   0.06209   0.05637  -0.0908   0.0015   1.0000
  13.750   1.6225   0.06622   0.06063  -0.0913   0.0015   1.0000
  14.000   1.6151   0.07071   0.06524  -0.0920   0.0015   1.0000
  14.250   1.6091   0.07520   0.06986  -0.0929   0.0015   1.0000
  14.500   1.6018   0.08003   0.07482  -0.0939   0.0014   1.0000
  14.750   1.5945   0.08497   0.07989  -0.0951   0.0014   1.0000
  15.000   1.5881   0.08992   0.08497  -0.0964   0.0014   1.0000
  15.250   1.5802   0.09526   0.09045  -0.0980   0.0013   1.0000
  15.500   1.5730   0.10055   0.09587  -0.0996   0.0013   1.0000
<< Back to EPPLER 377 AIRFOIL (e377-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 377 AIRFOIL (e377-il)