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

SOKOLOV AIRFOIL (sokolov-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: SOKOLOV AIRFOIL (sokolov-il)
Reynolds number: 50,000
Max Cl/Cd: 44.29 at α=8°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-sokolov-il-50000.txt
Download as CSV file: xf-sokolov-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: SOKOLOV 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.2838   0.11182   0.10515  -0.0305   1.0000   0.1219
  -7.750  -0.2869   0.11056   0.10400  -0.0302   1.0000   0.1254
  -7.500  -0.2999   0.11102   0.10464  -0.0294   1.0000   0.1272
  -7.250  -0.3124   0.11216   0.10597  -0.0308   1.0000   0.1280
  -7.000  -0.2932   0.10427   0.09802  -0.0267   1.0000   0.1329
  -6.750  -0.2950   0.10242   0.09627  -0.0255   1.0000   0.1371
  -6.500  -0.3032   0.10202   0.09601  -0.0255   1.0000   0.1406
  -6.250  -0.3131   0.10332   0.09746  -0.0292   1.0000   0.1422
  -6.000  -0.3073   0.09736   0.09156  -0.0228   1.0000   0.1465
  -5.750  -0.3097   0.09564   0.08992  -0.0215   1.0000   0.1521
  -5.500  -0.3127   0.09655   0.09092  -0.0275   1.0000   0.1566
  -5.250  -0.3140   0.09200   0.08645  -0.0210   1.0000   0.1610
  -5.000  -0.3124   0.09056   0.08507  -0.0221   1.0000   0.1691
  -4.750  -0.3095   0.08817   0.08274  -0.0231   1.0000   0.1738
  -4.500  -0.3064   0.08582   0.08044  -0.0219   1.0000   0.1813
  -4.000  -0.2764   0.08189   0.07650  -0.0327   1.0000   0.2009
  -3.750  -0.2791   0.07845   0.07312  -0.0272   1.0000   0.2053
  -3.500  -0.2617   0.07596   0.07062  -0.0313   1.0000   0.2178
  -3.250  -0.2467   0.07343   0.06808  -0.0335   1.0000   0.2324
  -3.000  -0.2341   0.07091   0.06557  -0.0340   1.0000   0.2489
  -2.750  -0.2081   0.06870   0.06330  -0.0393   1.0000   0.2748
  -2.000  -0.1567   0.06147   0.05604  -0.0421   1.0000   0.3498
  -0.750   0.2533   0.04117   0.03259  -0.1219   0.9769   0.1828
  -0.500   0.3161   0.03911   0.02980  -0.1296   0.9713   0.1751
  -0.250   0.3615   0.03836   0.02858  -0.1340   0.9641   0.1835
   0.000   0.4119   0.03761   0.02750  -0.1389   0.9574   0.1927
   0.250   0.4522   0.03731   0.02681  -0.1419   0.9498   0.2045
   0.500   0.4978   0.03699   0.02633  -0.1455   0.9430   0.2373
   0.750   0.5343   0.03672   0.02639  -0.1475   0.9354   0.3371
   1.000   0.5730   0.03490   0.02588  -0.1493   0.9285   1.0000
   1.250   0.6016   0.03586   0.02636  -0.1504   0.9179   1.0000
   1.500   0.6369   0.03662   0.02681  -0.1523   0.9055   1.0000
   1.750   0.6732   0.03728   0.02723  -0.1542   0.8922   1.0000
   2.000   0.7085   0.03793   0.02770  -0.1559   0.8794   1.0000
   2.250   0.7452   0.03855   0.02819  -0.1577   0.8680   1.0000
   2.500   0.7807   0.03920   0.02877  -0.1594   0.8572   1.0000
   2.750   0.8030   0.04023   0.02976  -0.1592   0.8453   1.0000
   3.000   0.8275   0.04126   0.03076  -0.1593   0.8337   1.0000
   3.250   0.8562   0.04215   0.03165  -0.1599   0.8224   1.0000
   3.500   0.8950   0.04264   0.03219  -0.1616   0.8122   1.0000
   3.750   0.9205   0.04360   0.03321  -0.1616   0.7999   1.0000
   4.000   0.9416   0.04477   0.03443  -0.1610   0.7868   1.0000
   4.250   0.9633   0.04596   0.03570  -0.1606   0.7737   1.0000
   4.500   0.9852   0.04719   0.03706  -0.1601   0.7606   1.0000
   4.750   1.0096   0.04824   0.03822  -0.1597   0.7470   1.0000
   5.000   1.0355   0.04910   0.03921  -0.1593   0.7327   1.0000
   5.250   1.0627   0.04978   0.04003  -0.1588   0.7180   1.0000
   5.500   1.0902   0.05036   0.04082  -0.1581   0.7034   1.0000
   5.750   1.1187   0.05078   0.04142  -0.1574   0.6887   1.0000
   6.000   1.1476   0.05104   0.04189  -0.1565   0.6742   1.0000
   6.250   1.1740   0.05145   0.04251  -0.1553   0.6598   1.0000
   6.500   1.1979   0.05209   0.04342  -0.1539   0.6456   1.0000
   6.750   1.2253   0.05220   0.04379  -0.1524   0.6311   1.0000
   7.000   1.2566   0.05173   0.04362  -0.1508   0.6162   1.0000
   7.250   1.2859   0.05148   0.04369  -0.1491   0.6015   1.0000
   7.500   1.3871   0.03905   0.03178  -0.1445   0.5652   1.0000
   7.750   1.4161   0.03314   0.02584  -0.1367   0.5031   1.0000
   8.000   1.4189   0.03204   0.02481  -0.1301   0.4400   1.0000
   8.250   1.4066   0.03280   0.02485  -0.1221   0.3265   1.0000
   8.500   1.3819   0.03636   0.02738  -0.1152   0.1991   1.0000
   8.750   1.3587   0.04031   0.03056  -0.1095   0.1617   1.0000
   9.000   1.3446   0.04419   0.03419  -0.1053   0.1357   1.0000
   9.250   1.3394   0.04743   0.03729  -0.1021   0.1173   1.0000
   9.500   1.3408   0.05018   0.04000  -0.0995   0.1047   1.0000
   9.750   1.3525   0.05234   0.04220  -0.0967   0.0955   1.0000
  10.000   1.3936   0.05373   0.04353  -0.0947   0.0863   1.0000
  10.250   1.4565   0.05669   0.04642  -0.0964   0.0768   1.0000
  10.500   1.5000   0.06108   0.05139  -0.0973   0.0743   1.0000
  10.750   1.5199   0.06558   0.05638  -0.0960   0.0739   1.0000
  11.000   1.5272   0.07011   0.06139  -0.0938   0.0742   1.0000
  11.250   1.5260   0.07460   0.06632  -0.0911   0.0748   1.0000
  11.500   1.5174   0.07887   0.07096  -0.0879   0.0755   1.0000
  11.750   1.5038   0.08308   0.07550  -0.0847   0.0760   1.0000
  12.000   1.4880   0.08751   0.08022  -0.0821   0.0765   1.0000
  12.250   1.4705   0.09222   0.08519  -0.0802   0.0770   1.0000
  12.500   1.4526   0.09730   0.09050  -0.0791   0.0775   1.0000
  12.750   1.4357   0.10276   0.09617  -0.0787   0.0781   1.0000
  13.000   1.4215   0.10868   0.10224  -0.0789   0.0786   1.0000
  13.250   1.3641   0.11556   0.10957  -0.0828   0.0802   1.0000
  13.500   1.3114   0.12628   0.12053  -0.0904   0.0815   1.0000
  13.750   1.2596   0.14103   0.13535  -0.1018   0.0841   1.0000
  14.000   1.2309   0.15451   0.14882  -0.1112   0.0873   1.0000
  14.250   1.2256   0.16270   0.15698  -0.1151   0.0896   1.0000
<< Back to SOKOLOV AIRFOIL (sokolov-il)

Polar data table (+)

Polar graphs


<< Back to SOKOLOV AIRFOIL (sokolov-il)