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

EPPLER 638 AIRFOIL (e638-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 638 AIRFOIL (e638-il)
Reynolds number: 50,000
Max Cl/Cd: 25.27 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e638-il-50000-n5.txt
Download as CSV file: xf-e638-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 638 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3033   0.10157   0.09557  -0.0239   1.0000   0.1087
  -8.500  -0.3110   0.09924   0.09338  -0.0265   1.0000   0.1132
  -8.250  -0.3333   0.09758   0.09191  -0.0307   1.0000   0.1148
  -8.000  -0.3103   0.09259   0.08699  -0.0280   1.0000   0.1190
  -7.750  -0.3076   0.08971   0.08423  -0.0279   1.0000   0.1239
  -7.500  -0.3195   0.08748   0.08219  -0.0286   1.0000   0.1273
  -7.250  -0.3532   0.08710   0.08197  -0.0308   1.0000   0.1306
  -7.000  -0.3182   0.08103   0.07601  -0.0316   0.9714   0.1360
  -6.750  -0.3058   0.07694   0.07177  -0.0403   0.9198   0.1479
  -6.500  -0.2733   0.07183   0.06660  -0.0428   0.8925   0.1536
  -6.000  -0.2307   0.05835   0.05168  -0.0538   0.8429   0.0564
  -5.750  -0.2126   0.05469   0.04789  -0.0539   0.8215   0.0549
  -5.500  -0.1978   0.05166   0.04457  -0.0533   0.8010   0.0538
  -5.250  -0.1828   0.04892   0.04146  -0.0524   0.7828   0.0538
  -5.000  -0.1671   0.04635   0.03844  -0.0511   0.7662   0.0544
  -4.750  -0.1501   0.04393   0.03558  -0.0497   0.7508   0.0548
  -4.500  -0.1320   0.04160   0.03282  -0.0482   0.7363   0.0546
  -4.250  -0.1125   0.03935   0.03013  -0.0467   0.7227   0.0542
  -4.000  -0.0914   0.03729   0.02760  -0.0453   0.7102   0.0542
  -3.750  -0.0685   0.03553   0.02528  -0.0438   0.6986   0.0556
  -3.500  -0.0457   0.03381   0.02319  -0.0427   0.6860   0.0579
  -3.250  -0.0213   0.03227   0.02142  -0.0419   0.6741   0.0599
  -3.000   0.0055   0.03086   0.01965  -0.0411   0.6636   0.0612
  -2.750   0.0334   0.02968   0.01816  -0.0405   0.6532   0.0644
  -2.500   0.0626   0.02873   0.01688  -0.0400   0.6423   0.0689
  -2.250   0.0959   0.02758   0.01554  -0.0403   0.6326   0.0721
  -2.000   0.1281   0.02673   0.01451  -0.0406   0.6230   0.0769
  -1.750   0.1565   0.02621   0.01377  -0.0403   0.6135   0.0851
  -1.500   0.1822   0.02555   0.01301  -0.0397   0.6058   0.0934
  -1.250   0.3280   0.02148   0.01164  -0.0597   0.5919   1.0000
  -1.000   0.3504   0.02169   0.01146  -0.0588   0.5836   1.0000
  -0.750   0.3729   0.02188   0.01130  -0.0578   0.5763   1.0000
  -0.500   0.3948   0.02214   0.01132  -0.0569   0.5684   1.0000
  -0.250   0.4170   0.02238   0.01130  -0.0560   0.5616   1.0000
   0.000   0.4388   0.02268   0.01138  -0.0551   0.5547   1.0000
   0.250   0.4605   0.02299   0.01151  -0.0542   0.5478   1.0000
   0.500   0.4828   0.02327   0.01156  -0.0532   0.5423   1.0000
   0.750   0.5034   0.02370   0.01191  -0.0522   0.5350   1.0000
   1.000   0.5256   0.02401   0.01203  -0.0512   0.5297   1.0000
   1.250   0.5462   0.02448   0.01241  -0.0503   0.5234   1.0000
   1.500   0.5671   0.02491   0.01275  -0.0493   0.5175   1.0000
   1.750   0.5898   0.02522   0.01288  -0.0483   0.5130   1.0000
   2.000   0.6085   0.02588   0.01355  -0.0473   0.5065   1.0000
   2.250   0.6294   0.02634   0.01393  -0.0462   0.5013   1.0000
   2.500   0.6523   0.02667   0.01412  -0.0453   0.4973   1.0000
   2.750   0.6690   0.02751   0.01502  -0.0441   0.4907   1.0000
   3.000   0.6895   0.02803   0.01550  -0.0430   0.4859   1.0000
   3.250   0.7125   0.02840   0.01576  -0.0421   0.4822   1.0000
   3.500   0.7271   0.02941   0.01688  -0.0408   0.4759   1.0000
   3.750   0.7465   0.03005   0.01750  -0.0397   0.4711   1.0000
   4.000   0.7692   0.03044   0.01782  -0.0388   0.4675   1.0000
   4.500   0.7990   0.03245   0.02000  -0.0360   0.4568   1.0000
   4.750   0.8209   0.03293   0.02044  -0.0351   0.4531   1.0000
   5.000   0.8337   0.03411   0.02170  -0.0336   0.4480   1.0000
   5.250   0.8454   0.03535   0.02307  -0.0320   0.4426   1.0000
   5.500   0.8655   0.03596   0.02368  -0.0309   0.4388   1.0000
   5.750   0.8821   0.03687   0.02463  -0.0297   0.4348   1.0000
   6.000   0.8829   0.03890   0.02682  -0.0274   0.4283   1.0000
   6.250   0.9004   0.03973   0.02769  -0.0263   0.4243   1.0000
   6.500   0.9260   0.04001   0.02799  -0.0256   0.4215   1.0000
   6.750   0.9042   0.04368   0.03187  -0.0223   0.4132   1.0000
   7.000   0.9215   0.04451   0.03276  -0.0211   0.4092   1.0000
   7.250   0.9488   0.04467   0.03294  -0.0206   0.4065   1.0000
   7.500   0.8986   0.05055   0.03897  -0.0165   0.3962   1.0000
   7.750   0.9233   0.05081   0.03928  -0.0157   0.3931   1.0000
   8.250   0.8783   0.05916   0.04771  -0.0122   0.3781   1.0000
   8.500   0.8355   0.06667   0.05523  -0.0130   0.3664   1.0000
   8.750   0.8569   0.06714   0.05580  -0.0121   0.3631   1.0000
   9.250   0.8454   0.07418   0.06293  -0.0121   0.3490   1.0000
   9.750   0.8390   0.08079   0.06965  -0.0122   0.3354   1.0000
  10.000   0.8651   0.08073   0.06968  -0.0111   0.3325   1.0000
  10.250   0.8376   0.08696   0.07594  -0.0125   0.3221   1.0000
  10.500   0.8607   0.08724   0.07632  -0.0116   0.3186   1.0000
  10.750   0.8406   0.09270   0.08182  -0.0128   0.3090   1.0000
  11.000   0.8593   0.09351   0.08277  -0.0121   0.3048   1.0000
  11.250   0.8462   0.09825   0.08756  -0.0132   0.2960   1.0000
  11.500   0.8621   0.09933   0.08875  -0.0126   0.2910   1.0000
<< Back to EPPLER 638 AIRFOIL (e638-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 638 AIRFOIL (e638-il)