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

EPPLER 542 AIRFOIL (e542-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 542 AIRFOIL (e542-il)
Reynolds number: 500,000
Max Cl/Cd: 79.32 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e542-il-500000.txt
Download as CSV file: xf-e542-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 542 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
  ------ -------- --------- --------- -------- -------- --------
 -13.250  -0.4119   0.12168   0.11935  -0.0477   1.0000   0.0153
 -13.000  -0.4114   0.11794   0.11562  -0.0488   1.0000   0.0154
  -9.500  -0.5975   0.02780   0.02243  -0.0678   0.9373   0.0093
  -9.250  -0.5557   0.02477   0.01912  -0.0716   0.9303   0.0090
  -9.000  -0.5152   0.02271   0.01662  -0.0748   0.9186   0.0096
  -8.750  -0.4693   0.02095   0.01466  -0.0785   0.9088   0.0097
  -8.500  -0.4305   0.01972   0.01327  -0.0807   0.8950   0.0097
  -8.250  -0.4053   0.01919   0.01268  -0.0809   0.8780   0.0105
  -8.000  -0.3813   0.01854   0.01190  -0.0805   0.8636   0.0110
  -7.750  -0.3596   0.01799   0.01123  -0.0797   0.8507   0.0114
  -7.500  -0.3388   0.01752   0.01063  -0.0786   0.8391   0.0118
  -7.250  -0.3211   0.01681   0.00984  -0.0771   0.8284   0.0121
  -7.000  -0.3062   0.01617   0.00914  -0.0751   0.8174   0.0123
  -6.750  -0.2910   0.01561   0.00853  -0.0733   0.8080   0.0129
  -6.500  -0.2751   0.01514   0.00800  -0.0715   0.7985   0.0138
  -6.250  -0.2597   0.01463   0.00742  -0.0695   0.7901   0.0153
  -6.000  -0.2433   0.01421   0.00694  -0.0677   0.7820   0.0178
  -5.750  -0.2312   0.01358   0.00635  -0.0652   0.7746   0.0303
  -5.500  -0.2201   0.01298   0.00590  -0.0625   0.7668   0.0604
  -5.250  -0.2097   0.01243   0.00551  -0.0599   0.7600   0.1039
  -5.000  -0.2055   0.01171   0.00510  -0.0561   0.7525   0.1711
  -4.750  -0.2103   0.01097   0.00466  -0.0507   0.7458   0.2522
  -4.500  -0.2231   0.01016   0.00428  -0.0435   0.7390   0.3520
  -4.250  -0.2381   0.00895   0.00368  -0.0361   0.7327   0.4990
  -4.000  -0.2383   0.00841   0.00401  -0.0304   0.7275   0.7035
  -3.750  -0.2121   0.00881   0.00436  -0.0298   0.7221   0.7343
  -3.500  -0.1842   0.00928   0.00475  -0.0294   0.7169   0.7483
  -3.250  -0.1556   0.00983   0.00518  -0.0292   0.7123   0.7595
  -3.000  -0.1237   0.01049   0.00585  -0.0293   0.7077   0.7648
  -2.750  -0.0919   0.01123   0.00655  -0.0294   0.7031   0.7714
  -2.500  -0.0602   0.01202   0.00728  -0.0293   0.6989   0.7789
  -2.250  -0.0072   0.01396   0.00919  -0.0316   0.6951   0.7828
  -2.000   0.0107   0.01408   0.00928  -0.0299   0.6904   0.7933
  -1.750   0.0415   0.01421   0.00936  -0.0303   0.6861   0.7942
  -1.500   0.0709   0.01427   0.00935  -0.0306   0.6824   0.7950
  -1.250   0.1001   0.01432   0.00933  -0.0309   0.6786   0.7959
  -1.000   0.1268   0.01430   0.00929  -0.0308   0.6744   0.7970
  -0.750   0.1533   0.01427   0.00922  -0.0308   0.6704   0.7984
  -0.500   0.1785   0.01419   0.00910  -0.0305   0.6667   0.8003
  -0.250   0.2023   0.01409   0.00891  -0.0303   0.6631   0.8029
   0.000   0.2147   0.01348   0.00828  -0.0287   0.6592   0.8089
   0.250   0.2404   0.01336   0.00814  -0.0285   0.6553   0.8097
   0.500   0.2673   0.01331   0.00807  -0.0285   0.6517   0.8104
   0.750   0.2948   0.01328   0.00800  -0.0287   0.6482   0.8111
   1.000   0.3209   0.01323   0.00793  -0.0287   0.6447   0.8119
   1.250   0.3465   0.01316   0.00788  -0.0285   0.6406   0.8128
   1.500   0.3725   0.01311   0.00782  -0.0284   0.6366   0.8139
   1.750   0.3987   0.01303   0.00771  -0.0284   0.6328   0.8148
   2.000   0.4244   0.01295   0.00761  -0.0285   0.6285   0.8158
   2.250   0.4490   0.01283   0.00751  -0.0282   0.6239   0.8172
   2.500   0.4747   0.01272   0.00740  -0.0283   0.6195   0.8186
   2.750   0.5017   0.01263   0.00726  -0.0286   0.6155   0.8197
   3.000   0.5271   0.01250   0.00715  -0.0287   0.6110   0.8209
   3.250   0.5530   0.01236   0.00702  -0.0289   0.6060   0.8221
   3.500   0.5803   0.01228   0.00692  -0.0294   0.6013   0.8235
   3.750   0.6079   0.01224   0.00687  -0.0299   0.5966   0.8244
   4.000   0.6323   0.01209   0.00677  -0.0295   0.5911   0.8250
   4.250   0.6578   0.01200   0.00670  -0.0294   0.5858   0.8256
   4.500   0.6839   0.01197   0.00667  -0.0294   0.5805   0.8262
   4.750   0.7081   0.01189   0.00665  -0.0290   0.5741   0.8268
   5.000   0.7337   0.01186   0.00663  -0.0289   0.5681   0.8274
   5.250   0.7581   0.01183   0.00665  -0.0285   0.5615   0.8281
   5.500   0.7822   0.01180   0.00667  -0.0280   0.5543   0.8290
   5.750   0.8063   0.01180   0.00670  -0.0276   0.5469   0.8299
   6.000   0.8299   0.01177   0.00671  -0.0271   0.5382   0.8306
   6.250   0.8532   0.01175   0.00674  -0.0266   0.5287   0.8314
   6.500   0.8756   0.01177   0.00675  -0.0259   0.5179   0.8322
   6.750   0.8974   0.01177   0.00680  -0.0251   0.5054   0.8331
   7.000   0.9183   0.01181   0.00685  -0.0241   0.4909   0.8340
   7.250   0.9369   0.01189   0.00693  -0.0227   0.4728   0.8350
   7.500   0.9542   0.01203   0.00703  -0.0211   0.4516   0.8360
   7.750   0.9658   0.01224   0.00717  -0.0184   0.4263   0.8370
   8.000   0.9730   0.01256   0.00738  -0.0149   0.3964   0.8383
   8.250   0.9797   0.01304   0.00774  -0.0115   0.3661   0.8396
   8.500   0.9839   0.01363   0.00818  -0.0078   0.3341   0.8406
   8.750   0.9870   0.01427   0.00870  -0.0040   0.3034   0.8417
   9.000   0.9909   0.01496   0.00929  -0.0006   0.2754   0.8428
   9.250   0.9953   0.01570   0.00995   0.0026   0.2512   0.8439
   9.500   1.0005   0.01648   0.01065   0.0054   0.2281   0.8451
  10.000   1.0105   0.01824   0.01227   0.0107   0.1869   0.8474
  10.250   1.0159   0.01921   0.01318   0.0130   0.1684   0.8487
  10.500   1.0203   0.02029   0.01420   0.0152   0.1505   0.8499
  10.750   1.0252   0.02142   0.01527   0.0173   0.1344   0.8513
  11.000   1.0307   0.02258   0.01639   0.0191   0.1192   0.8527
  11.250   1.0361   0.02380   0.01757   0.0208   0.1052   0.8540
  11.500   1.0415   0.02509   0.01881   0.0223   0.0923   0.8550
  11.750   1.0457   0.02647   0.02013   0.0239   0.0786   0.8560
  12.000   1.0502   0.02782   0.02148   0.0254   0.0686   0.8572
  12.250   1.0543   0.02928   0.02292   0.0268   0.0589   0.8583
  12.500   1.0583   0.03080   0.02442   0.0281   0.0503   0.8595
  12.750   1.0623   0.03238   0.02600   0.0293   0.0442   0.8607
  13.000   1.0672   0.03396   0.02759   0.0303   0.0382   0.8619
  13.250   1.0728   0.03554   0.02919   0.0311   0.0328   0.8632
  13.500   1.0780   0.03721   0.03089   0.0318   0.0292   0.8645
  13.750   1.0829   0.03897   0.03265   0.0324   0.0256   0.8658
  14.000   1.0893   0.04067   0.03442   0.0329   0.0233   0.8671
  14.250   1.0941   0.04255   0.03633   0.0333   0.0209   0.8684
  14.500   1.0998   0.04441   0.03825   0.0336   0.0191   0.8697
  14.750   1.1028   0.04658   0.04047   0.0339   0.0176   0.8710
  15.000   1.1089   0.04842   0.04240   0.0340   0.0158   0.8726
  15.250   1.1092   0.05092   0.04496   0.0342   0.0147   0.8740
  15.500   1.1157   0.05285   0.04697   0.0342   0.0128   0.8755
  15.750   1.1161   0.05550   0.04969   0.0341   0.0118   0.8769
  16.000   1.1190   0.05796   0.05225   0.0339   0.0102   0.8784
  16.250   1.1163   0.06114   0.05550   0.0335   0.0095   0.8798
  16.500   1.1196   0.06370   0.05817   0.0330   0.0081   0.8814
  16.750   1.1174   0.06702   0.06155   0.0323   0.0072   0.8829
  17.000   1.1162   0.07036   0.06500   0.0315   0.0061   0.8844
  17.250   1.1117   0.07423   0.06895   0.0304   0.0057   0.8857
  17.500   1.1078   0.07808   0.07294   0.0292   0.0053   0.8872
  17.750   1.1025   0.08221   0.07717   0.0278   0.0043   0.8888
  18.000   1.0994   0.08616   0.08125   0.0264   0.0044   0.8905
  18.250   1.0944   0.09050   0.08571   0.0246   0.0043   0.8921
  18.500   1.0857   0.09552   0.09087   0.0225   0.0043   0.8937
  18.750   1.0732   0.10136   0.09685   0.0199   0.0038   0.8952
<< Back to EPPLER 542 AIRFOIL (e542-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 542 AIRFOIL (e542-il)