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

EH 0.0/9.0 (eh0009-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EH 0.0/9.0 (eh0009-il)
Reynolds number: 50,000
Max Cl/Cd: 26.66 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-eh0009-il-50000.txt
Download as CSV file: xf-eh0009-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EH 0.0/9.0                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7324   0.08076   0.07436  -0.0161   1.0000   0.1176
  -8.750  -0.7284   0.07567   0.06914  -0.0165   1.0000   0.1153
  -8.500  -0.7605   0.06814   0.06130  -0.0194   1.0000   0.1034
  -8.250  -0.7599   0.06316   0.05614  -0.0193   1.0000   0.1027
  -8.000  -0.7599   0.05833   0.05101  -0.0190   1.0000   0.1022
  -7.750  -0.7584   0.05363   0.04588  -0.0182   1.0000   0.1024
  -7.500  -0.7525   0.04916   0.04080  -0.0172   1.0000   0.1027
  -7.250  -0.7417   0.04484   0.03593  -0.0158   1.0000   0.1029
  -7.000  -0.7274   0.04087   0.03129  -0.0143   1.0000   0.1038
  -6.750  -0.7069   0.03731   0.02773  -0.0133   1.0000   0.1096
  -6.500  -0.6883   0.03455   0.02459  -0.0120   1.0000   0.1208
  -6.250  -0.6658   0.03165   0.02129  -0.0106   1.0000   0.1318
  -6.000  -0.6425   0.02913   0.01881  -0.0094   1.0000   0.1510
  -5.750  -0.6210   0.02690   0.01654  -0.0078   1.0000   0.1847
  -5.500  -0.6018   0.02467   0.01474  -0.0059   1.0000   0.2404
  -5.250  -0.5877   0.02298   0.01363  -0.0032   1.0000   0.3233
  -5.000  -0.5739   0.02172   0.01293   0.0002   1.0000   0.4133
  -4.750  -0.5592   0.02097   0.01251   0.0040   1.0000   0.4964
  -4.500  -0.5436   0.02053   0.01232   0.0081   1.0000   0.5694
  -4.250  -0.5265   0.02036   0.01232   0.0125   1.0000   0.6317
  -4.000  -0.5089   0.02032   0.01234   0.0169   1.0000   0.6880
  -3.750  -0.4895   0.02041   0.01241   0.0213   1.0000   0.7388
  -3.500  -0.4664   0.02058   0.01247   0.0251   1.0000   0.7871
  -3.250  -0.4288   0.02095   0.01256   0.0265   1.0000   0.8335
  -3.000  -0.3600   0.02149   0.01264   0.0219   1.0000   0.8796
  -2.750  -0.2623   0.02157   0.01210   0.0101   1.0000   0.9223
  -2.500  -0.1727   0.02085   0.01096  -0.0020   1.0000   0.9601
  -2.250  -0.0851   0.01956   0.00931  -0.0152   1.0000   0.9947
  -2.000  -0.0606   0.01875   0.00841  -0.0170   1.0000   1.0000
  -1.750  -0.0492   0.01818   0.00781  -0.0162   1.0000   1.0000
  -1.500  -0.0382   0.01770   0.00731  -0.0150   1.0000   1.0000
  -1.250  -0.0279   0.01729   0.00690  -0.0135   1.0000   1.0000
  -1.000  -0.0188   0.01697   0.00659  -0.0116   1.0000   1.0000
  -0.750  -0.0115   0.01671   0.00634  -0.0093   1.0000   1.0000
  -0.500  -0.0062   0.01653   0.00618  -0.0065   1.0000   1.0000
  -0.250  -0.0024   0.01642   0.00609  -0.0034   1.0000   1.0000
   0.000   0.0000   0.01638   0.00605   0.0000   1.0000   1.0000
   0.250   0.0025   0.01642   0.00608   0.0034   1.0000   1.0000
   0.500   0.0062   0.01653   0.00618   0.0065   1.0000   1.0000
   0.750   0.0115   0.01671   0.00634   0.0093   1.0000   1.0000
   1.000   0.0189   0.01697   0.00658   0.0116   1.0000   1.0000
   1.250   0.0280   0.01729   0.00690   0.0135   1.0000   1.0000
   1.500   0.0383   0.01769   0.00731   0.0150   1.0000   1.0000
   1.750   0.0493   0.01817   0.00780   0.0162   1.0000   1.0000
   2.000   0.0607   0.01874   0.00840   0.0170   1.0000   1.0000
   2.250   0.0849   0.01955   0.00929   0.0152   0.9948   1.0000
   2.500   0.1728   0.02084   0.01095   0.0020   0.9602   1.0000
   2.750   0.2623   0.02156   0.01210  -0.0101   0.9224   1.0000
   3.000   0.3601   0.02148   0.01263  -0.0219   0.8796   1.0000
   3.250   0.4286   0.02095   0.01255  -0.0265   0.8337   1.0000
   3.500   0.4663   0.02057   0.01246  -0.0251   0.7873   1.0000
   3.750   0.4894   0.02041   0.01241  -0.0213   0.7389   1.0000
   4.000   0.5088   0.02032   0.01234  -0.0169   0.6881   1.0000
   4.250   0.5264   0.02036   0.01232  -0.0124   0.6319   1.0000
   4.500   0.5434   0.02053   0.01232  -0.0081   0.5695   1.0000
   4.750   0.5591   0.02097   0.01251  -0.0040   0.4966   1.0000
   5.000   0.5738   0.02172   0.01293  -0.0002   0.4135   1.0000
   5.250   0.5876   0.02298   0.01363   0.0032   0.3232   1.0000
   5.500   0.6018   0.02467   0.01474   0.0059   0.2404   1.0000
   5.750   0.6210   0.02690   0.01654   0.0078   0.1847   1.0000
   6.000   0.6424   0.02912   0.01880   0.0094   0.1508   1.0000
   6.250   0.6658   0.03164   0.02128   0.0106   0.1319   1.0000
   6.500   0.6883   0.03453   0.02456   0.0120   0.1207   1.0000
   6.750   0.7069   0.03730   0.02771   0.0133   0.1094   1.0000
   7.000   0.7274   0.04089   0.03130   0.0143   0.1038   1.0000
   7.250   0.7417   0.04484   0.03593   0.0158   0.1029   1.0000
   7.500   0.7526   0.04916   0.04080   0.0171   0.1026   1.0000
   7.750   0.7584   0.05366   0.04592   0.0182   0.1024   1.0000
   8.000   0.7603   0.05834   0.05102   0.0189   0.1023   1.0000
   8.250   0.7595   0.06320   0.05620   0.0192   0.1026   1.0000
   8.500   0.7609   0.06819   0.06134   0.0193   0.1034   1.0000
   8.750   0.7283   0.07572   0.06919   0.0163   0.1153   1.0000
   9.000   0.7332   0.08083   0.07442   0.0160   0.1176   1.0000
   9.250   0.5804   0.08585   0.07963   0.0060   0.1411   1.0000
<< Back to EH 0.0/9.0 (eh0009-il)

Polar data table (+)

Polar graphs


<< Back to EH 0.0/9.0 (eh0009-il)