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

EH 2.0/7.0 (eh2070-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EH 2.0/7.0 (eh2070-il)
Reynolds number: 50,000
Max Cl/Cd: 33.58 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-eh2070-il-50000-n5.txt
Download as CSV file: xf-eh2070-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 2.0/7.0                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5390   0.09643   0.09023   0.0077   1.0000   0.1025
  -7.500  -0.5420   0.09318   0.08706   0.0044   1.0000   0.1065
  -7.250  -0.5523   0.09103   0.08487  -0.0036   1.0000   0.1092
  -7.000  -0.5359   0.08524   0.07920  -0.0007   1.0000   0.1121
  -6.750  -0.5277   0.08127   0.07524  -0.0017   1.0000   0.1150
  -6.500  -0.5280   0.07834   0.07218  -0.0074   1.0000   0.1231
  -6.000  -0.4881   0.06527   0.05875  -0.0113   1.0000   0.0528
  -5.750  -0.4750   0.06082   0.05418  -0.0123   1.0000   0.0487
  -5.500  -0.4552   0.05599   0.04879  -0.0140   1.0000   0.0425
  -5.250  -0.4400   0.05194   0.04463  -0.0141   1.0000   0.0414
  -5.000  -0.4227   0.04812   0.04058  -0.0141   1.0000   0.0405
  -4.750  -0.4037   0.04442   0.03657  -0.0139   1.0000   0.0397
  -4.500  -0.3832   0.04090   0.03268  -0.0135   1.0000   0.0391
  -4.250  -0.3613   0.03761   0.02897  -0.0129   1.0000   0.0387
  -4.000  -0.3383   0.03455   0.02546  -0.0122   1.0000   0.0385
  -3.750  -0.3141   0.03173   0.02218  -0.0113   1.0000   0.0388
  -3.500  -0.2887   0.02925   0.01922  -0.0105   1.0000   0.0401
  -3.250  -0.2629   0.02713   0.01667  -0.0097   1.0000   0.0444
  -3.000  -0.2382   0.02548   0.01475  -0.0091   1.0000   0.0498
  -2.750  -0.2104   0.02363   0.01259  -0.0082   1.0000   0.0535
  -2.500  -0.1820   0.02195   0.01074  -0.0075   1.0000   0.0592
  -2.250  -0.1562   0.02083   0.00957  -0.0070   1.0000   0.0764
  -2.000  -0.1344   0.01964   0.00837  -0.0057   1.0000   0.0995
  -1.750  -0.0272   0.01496   0.00656  -0.0182   1.0000   1.0000
  -1.500  -0.0078   0.01491   0.00610  -0.0170   1.0000   1.0000
  -1.250   0.0093   0.01492   0.00589  -0.0157   1.0000   1.0000
  -1.000   0.0341   0.01500   0.00574  -0.0161   0.9826   1.0000
  -0.750   0.0822   0.01508   0.00548  -0.0206   0.9365   1.0000
  -0.500   0.1268   0.01516   0.00526  -0.0240   0.8966   1.0000
  -0.250   0.1626   0.01528   0.00506  -0.0254   0.8604   1.0000
   0.000   0.1900   0.01545   0.00499  -0.0251   0.8277   1.0000
   0.250   0.2141   0.01565   0.00497  -0.0240   0.7996   1.0000
   0.500   0.2367   0.01587   0.00499  -0.0227   0.7743   1.0000
   0.750   0.2590   0.01610   0.00505  -0.0213   0.7518   1.0000
   1.000   0.2813   0.01635   0.00515  -0.0201   0.7303   1.0000
   1.250   0.3038   0.01660   0.00526  -0.0189   0.7107   1.0000
   1.500   0.3266   0.01686   0.00542  -0.0178   0.6916   1.0000
   1.750   0.3494   0.01713   0.00559  -0.0166   0.6743   1.0000
   2.000   0.3726   0.01741   0.00582  -0.0157   0.6568   1.0000
   2.250   0.3960   0.01771   0.00615  -0.0147   0.6400   1.0000
   2.500   0.4195   0.01801   0.00643  -0.0138   0.6240   1.0000
   2.750   0.4432   0.01833   0.00674  -0.0129   0.6086   1.0000
   3.000   0.4669   0.01865   0.00707  -0.0121   0.5937   1.0000
   3.250   0.4908   0.01899   0.00745  -0.0112   0.5789   1.0000
   3.500   0.5147   0.01935   0.00789  -0.0105   0.5638   1.0000
   3.750   0.5389   0.01973   0.00849  -0.0098   0.5489   1.0000
   4.000   0.5632   0.02013   0.00901  -0.0091   0.5341   1.0000
   4.250   0.5875   0.02054   0.00956  -0.0084   0.5195   1.0000
   4.500   0.6116   0.02097   0.01015  -0.0077   0.5049   1.0000
   4.750   0.6357   0.02141   0.01077  -0.0070   0.4902   1.0000
   5.000   0.6598   0.02188   0.01143  -0.0062   0.4756   1.0000
   5.250   0.6839   0.02236   0.01215  -0.0054   0.4610   1.0000
   5.500   0.7079   0.02286   0.01307  -0.0045   0.4462   1.0000
   5.750   0.7318   0.02339   0.01392  -0.0037   0.4300   1.0000
   6.000   0.7553   0.02389   0.01484  -0.0027   0.4116   1.0000
   6.250   0.7693   0.02291   0.01367   0.0005   0.3288   1.0000
   6.500   0.7775   0.02458   0.01410   0.0022   0.1219   1.0000
   6.750   0.7865   0.02800   0.01684   0.0033   0.0607   1.0000
   7.000   0.7995   0.03031   0.01920   0.0044   0.0459   1.0000
   7.250   0.8107   0.03263   0.02159   0.0055   0.0388   1.0000
   7.500   0.8242   0.03459   0.02385   0.0070   0.0359   1.0000
   7.750   0.8378   0.03665   0.02615   0.0087   0.0335   1.0000
   8.000   0.8531   0.03883   0.02854   0.0103   0.0316   1.0000
   8.250   0.8704   0.04123   0.03129   0.0119   0.0302   1.0000
   8.500   0.8875   0.04416   0.03441   0.0131   0.0289   1.0000
   8.750   0.9019   0.04768   0.03818   0.0141   0.0276   1.0000
   9.000   0.9112   0.05058   0.04160   0.0151   0.0263   1.0000
   9.250   0.9160   0.05387   0.04533   0.0159   0.0252   1.0000
   9.500   0.9157   0.05741   0.04925   0.0164   0.0243   1.0000
   9.750   0.9105   0.06103   0.05318   0.0168   0.0241   1.0000
  10.000   0.8995   0.06478   0.05716   0.0167   0.0239   1.0000
  10.250   0.8872   0.06913   0.06171   0.0151   0.0241   1.0000
  10.500   0.8734   0.07415   0.06690   0.0122   0.0242   1.0000
  10.750   0.8596   0.07982   0.07271   0.0085   0.0246   1.0000
  11.000   0.8451   0.08626   0.07925   0.0041   0.0250   1.0000
  11.250   0.8325   0.09289   0.08596  -0.0002   0.0256   1.0000
<< Back to EH 2.0/7.0 (eh2070-il)

Polar data table (+)

Polar graphs


<< Back to EH 2.0/7.0 (eh2070-il)