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

EPPLER 376 AIRFOIL (e376-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 376 AIRFOIL (e376-il)
Reynolds number: 500,000
Max Cl/Cd: 134.98 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e376-il-500000.txt
Download as CSV file: xf-e376-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 376 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -1.250   0.4602   0.03307   0.02873  -0.0991   0.5722   0.0114
  -1.000   0.4939   0.03143   0.02700  -0.1014   0.5673   0.0126
  -0.750   0.5283   0.02984   0.02532  -0.1036   0.5619   0.0133
  -0.500   0.5633   0.02858   0.02392  -0.1055   0.5572   0.0141
  -0.250   0.6006   0.02781   0.02304  -0.1073   0.5523   0.0146
   0.000   0.6348   0.02668   0.02182  -0.1088   0.5475   0.0148
   0.250   0.6675   0.02542   0.02040  -0.1100   0.5430   0.0149
   0.500   0.6991   0.02288   0.01778  -0.1118   0.5386   0.0156
   0.750   0.7278   0.02155   0.01640  -0.1128   0.5337   0.0167
   1.000   0.7584   0.02055   0.01527  -0.1136   0.5295   0.0181
   1.250   0.7902   0.01960   0.01421  -0.1142   0.5251   0.0205
   1.500   0.8241   0.01942   0.01391  -0.1142   0.5203   0.0228
   2.500   0.9441   0.01519   0.00920  -0.1157   0.5028   0.0355
   2.750   0.9720   0.01462   0.00852  -0.1158   0.4986   0.0374
   3.000   1.0009   0.01397   0.00781  -0.1158   0.4940   0.0393
   3.250   1.0334   0.01179   0.00522  -0.1153   0.4900   0.0224
   3.500   1.0610   0.01086   0.00400  -0.1150   0.4859   0.0215
   3.750   1.0880   0.01049   0.00356  -0.1147   0.4812   0.0229
   4.250   1.1414   0.01040   0.00343  -0.1142   0.4706   0.0653
   4.500   1.1682   0.01049   0.00359  -0.1140   0.4645   0.0812
   4.750   1.1948   0.01052   0.00361  -0.1138   0.4591   0.0898
   5.000   1.2214   0.01057   0.00373  -0.1136   0.4536   0.1005
   5.250   1.2479   0.01062   0.00381  -0.1134   0.4472   0.1180
   5.500   1.2743   0.01072   0.00395  -0.1132   0.4408   0.1401
   5.750   1.3007   0.01079   0.00409  -0.1131   0.4339   0.1626
   6.000   1.3268   0.01092   0.00424  -0.1128   0.4275   0.1848
   6.250   1.3532   0.01097   0.00444  -0.1127   0.4199   0.2152
   6.500   1.3804   0.01028   0.00465  -0.1129   0.4123   1.0000
   6.750   1.4063   0.01044   0.00481  -0.1126   0.4034   1.0000
   7.000   1.4321   0.01061   0.00501  -0.1124   0.3942   1.0000
   7.250   1.4575   0.01081   0.00522  -0.1121   0.3843   1.0000
   7.500   1.4826   0.01103   0.00544  -0.1118   0.3668   1.0000
   7.750   1.5059   0.01149   0.00578  -0.1114   0.3293   1.0000
   8.000   1.5283   0.01213   0.00624  -0.1109   0.2883   1.0000
   8.250   1.5484   0.01314   0.00696  -0.1102   0.2351   1.0000
   8.500   1.5649   0.01463   0.00804  -0.1093   0.1697   1.0000
   8.750   1.5805   0.01617   0.00921  -0.1083   0.1141   1.0000
   9.000   1.5937   0.01789   0.01059  -0.1069   0.0623   1.0000
   9.250   1.6013   0.02012   0.01245  -0.1049   0.0141   1.0000
   9.500   1.6159   0.02137   0.01372  -0.1034   0.0078   1.0000
   9.750   1.6318   0.02235   0.01485  -0.1020   0.0070   1.0000
  10.000   1.6452   0.02348   0.01612  -0.1004   0.0064   1.0000
  10.250   1.6553   0.02478   0.01756  -0.0984   0.0060   1.0000
  10.500   1.6592   0.02629   0.01920  -0.0957   0.0056   1.0000
  10.750   1.6564   0.02800   0.02106  -0.0924   0.0054   1.0000
  11.000   1.6499   0.03037   0.02361  -0.0897   0.0052   1.0000
  11.250   1.6451   0.03304   0.02643  -0.0881   0.0051   1.0000
  11.500   1.6380   0.03631   0.02986  -0.0869   0.0050   1.0000
  11.750   1.6302   0.03997   0.03367  -0.0862   0.0049   1.0000
  12.000   1.6228   0.04378   0.03763  -0.0858   0.0049   1.0000
  12.250   1.6174   0.04744   0.04142  -0.0855   0.0049   1.0000
  12.500   1.6094   0.05152   0.04565  -0.0853   0.0049   1.0000
  12.750   1.6039   0.05534   0.04960  -0.0852   0.0049   1.0000
  13.000   1.5984   0.05920   0.05358  -0.0851   0.0049   1.0000
  13.250   1.5932   0.06308   0.05760  -0.0849   0.0049   1.0000
  13.500   1.5894   0.06684   0.06148  -0.0849   0.0050   1.0000
  13.750   1.5860   0.07065   0.06543  -0.0848   0.0050   1.0000
  14.000   1.5838   0.07441   0.06934  -0.0849   0.0051   1.0000
  14.250   1.5810   0.07835   0.07346  -0.0849   0.0053   1.0000
  14.500   1.5776   0.08257   0.07790  -0.0846   0.0055   1.0000
  14.750   1.5713   0.08747   0.08307  -0.0843   0.0059   1.0000
  15.000   1.5613   0.09332   0.08920  -0.0847   0.0063   1.0000
  15.250   1.5487   0.09977   0.09591  -0.0862   0.0067   1.0000
  15.500   1.5344   0.10667   0.10303  -0.0885   0.0069   1.0000
  15.750   1.5192   0.11399   0.11057  -0.0915   0.0071   1.0000
<< Back to EPPLER 376 AIRFOIL (e376-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 376 AIRFOIL (e376-il)