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

GOE 225 (MVA H.35) AIRFOIL (goe225-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 225 (MVA H.35) AIRFOIL (goe225-il)
Reynolds number: 200,000
Max Cl/Cd: 83.26 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe225-il-200000-n5.txt
Download as CSV file: xf-goe225-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 225 (MVA H.35) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.0004   0.08429   0.07999  -0.1069   0.8828   0.0382
  -7.750   0.0037   0.02764   0.02199  -0.1885   0.8614   0.0376
  -7.500   0.0419   0.02491   0.01881  -0.1944   0.8588   0.0382
  -7.250   0.0757   0.02319   0.01677  -0.1977   0.8557   0.0388
  -7.000   0.1077   0.02193   0.01525  -0.2000   0.8520   0.0395
  -6.750   0.1396   0.02090   0.01398  -0.2017   0.8481   0.0402
  -6.500   0.1705   0.02002   0.01303  -0.2030   0.8444   0.0411
  -6.250   0.2021   0.01928   0.01219  -0.2041   0.8412   0.0418
  -6.000   0.2327   0.01865   0.01146  -0.2051   0.8375   0.0426
  -5.750   0.2629   0.01809   0.01082  -0.2060   0.8337   0.0434
  -5.500   0.2937   0.01756   0.01020  -0.2069   0.8300   0.0444
  -5.250   0.3249   0.01707   0.00959  -0.2077   0.8268   0.0455
  -5.000   0.3564   0.01659   0.00904  -0.2086   0.8240   0.0467
  -4.750   0.3873   0.01622   0.00863  -0.2094   0.8208   0.0485
  -4.500   0.4168   0.01593   0.00831  -0.2099   0.8162   0.0511
  -4.250   0.4471   0.01561   0.00797  -0.2104   0.8118   0.0544
  -4.000   0.4779   0.01531   0.00763  -0.2110   0.8079   0.0591
  -3.750   0.5091   0.01504   0.00733  -0.2116   0.8047   0.0669
  -3.500   0.5375   0.01493   0.00723  -0.2117   0.7992   0.0776
  -3.250   0.5666   0.01482   0.00707  -0.2118   0.7938   0.0878
  -3.000   0.5966   0.01464   0.00686  -0.2121   0.7891   0.0969
  -2.750   0.6258   0.01451   0.00670  -0.2124   0.7839   0.1044
  -2.500   0.6542   0.01440   0.00657  -0.2124   0.7772   0.1122
  -2.250   0.6841   0.01420   0.00633  -0.2127   0.7715   0.1194
  -2.000   0.7126   0.01407   0.00621  -0.2128   0.7640   0.1272
  -1.750   0.7416   0.01389   0.00605  -0.2130   0.7563   0.1373
  -1.500   0.7704   0.01375   0.00591  -0.2131   0.7482   0.1485
  -1.250   0.7993   0.01357   0.00577  -0.2132   0.7386   0.1623
  -1.000   0.8274   0.01343   0.00572  -0.2133   0.7272   0.1875
  -0.750   0.8556   0.01331   0.00568  -0.2132   0.7146   0.2297
  -0.500   0.8835   0.01327   0.00562  -0.2130   0.6999   0.2613
  -0.250   0.9111   0.01324   0.00552  -0.2127   0.6823   0.2821
   0.000   0.9384   0.01325   0.00541  -0.2124   0.6628   0.2979
   0.250   0.9652   0.01331   0.00535  -0.2119   0.6435   0.3113
   0.500   0.9914   0.01344   0.00537  -0.2114   0.6261   0.3254
   0.750   1.0173   0.01362   0.00545  -0.2109   0.6105   0.3405
   1.000   1.0432   0.01383   0.00555  -0.2105   0.5969   0.3534
   1.250   1.0690   0.01404   0.00571  -0.2100   0.5855   0.3666
   1.500   1.0949   0.01426   0.00589  -0.2096   0.5753   0.3795
   1.750   1.1210   0.01448   0.00608  -0.2092   0.5663   0.3942
   2.000   1.1468   0.01472   0.00630  -0.2088   0.5578   0.4113
   2.250   1.1727   0.01496   0.00653  -0.2085   0.5504   0.4292
   2.500   1.1985   0.01519   0.00677  -0.2081   0.5430   0.4473
   2.750   1.2241   0.01547   0.00702  -0.2076   0.5364   0.4662
   3.000   1.2498   0.01568   0.00727  -0.2072   0.5296   0.4846
   3.250   1.2748   0.01595   0.00753  -0.2067   0.5220   0.5009
   3.500   1.2993   0.01621   0.00780  -0.2061   0.5136   0.5166
   3.750   1.3232   0.01649   0.00806  -0.2053   0.5046   0.5322
   4.000   1.3474   0.01676   0.00834  -0.2047   0.4972   0.5483
   4.250   1.3713   0.01699   0.00862  -0.2040   0.4901   0.5653
   4.750   1.4185   0.01749   0.00920  -0.2024   0.4774   0.6011
   5.000   1.4416   0.01772   0.00950  -0.2016   0.4708   0.6223
   5.250   1.4643   0.01798   0.00981  -0.2007   0.4650   0.6532
   5.500   1.4812   0.01779   0.01005  -0.1985   0.4589   0.9232
   5.750   1.5033   0.01811   0.01037  -0.1975   0.4525   1.0000
   6.000   1.5248   0.01848   0.01072  -0.1964   0.4462   1.0000
   6.250   1.5457   0.01881   0.01110  -0.1952   0.4389   1.0000
   6.500   1.5652   0.01921   0.01148  -0.1938   0.4322   1.0000
   6.750   1.5850   0.01958   0.01191  -0.1924   0.4254   1.0000
   7.000   1.6024   0.01998   0.01233  -0.1906   0.4182   1.0000
   7.250   1.6184   0.02043   0.01280  -0.1885   0.4110   1.0000
   7.500   1.6335   0.02092   0.01332  -0.1864   0.4017   1.0000
   7.750   1.6479   0.02147   0.01390  -0.1842   0.3918   1.0000
   8.000   1.6598   0.02215   0.01454  -0.1817   0.3805   1.0000
   8.250   1.6725   0.02283   0.01524  -0.1795   0.3678   1.0000
   8.500   1.6848   0.02358   0.01601  -0.1772   0.3554   1.0000
   8.750   1.6959   0.02444   0.01686  -0.1748   0.3432   1.0000
   9.000   1.7062   0.02538   0.01779  -0.1725   0.3312   1.0000
   9.250   1.7165   0.02637   0.01877  -0.1702   0.3184   1.0000
   9.500   1.7263   0.02743   0.01984  -0.1679   0.3057   1.0000
   9.750   1.7348   0.02861   0.02102  -0.1656   0.2933   1.0000
  10.000   1.7421   0.02991   0.02230  -0.1632   0.2814   1.0000
  10.250   1.7486   0.03131   0.02369  -0.1608   0.2696   1.0000
  10.500   1.7563   0.03267   0.02507  -0.1586   0.2587   1.0000
  10.750   1.7620   0.03421   0.02660  -0.1563   0.2484   1.0000
  11.000   1.7669   0.03584   0.02824  -0.1541   0.2386   1.0000
  11.250   1.7735   0.03738   0.02982  -0.1520   0.2300   1.0000
  11.500   1.7761   0.03927   0.03170  -0.1497   0.2212   1.0000
  11.750   1.7815   0.04096   0.03345  -0.1477   0.2119   1.0000
  12.000   1.7834   0.04300   0.03551  -0.1456   0.2034   1.0000
  12.250   1.7867   0.04500   0.03756  -0.1437   0.1942   1.0000
  12.500   1.7891   0.04714   0.03974  -0.1418   0.1860   1.0000
  12.750   1.7902   0.04946   0.04209  -0.1400   0.1776   1.0000
  13.000   1.7926   0.05172   0.04442  -0.1384   0.1693   1.0000
  13.250   1.7906   0.05452   0.04725  -0.1367   0.1599   1.0000
  13.500   1.7889   0.05741   0.05016  -0.1352   0.1485   1.0000
  13.750   1.7860   0.06050   0.05330  -0.1338   0.1364   1.0000
  14.000   1.7792   0.06415   0.05695  -0.1326   0.1218   1.0000
  14.250   1.7675   0.06853   0.06129  -0.1314   0.1021   1.0000
  14.500   1.7438   0.07462   0.06721  -0.1304   0.0750   1.0000
  14.750   1.7274   0.07992   0.07248  -0.1298   0.0654   1.0000
  15.000   1.7160   0.08469   0.07729  -0.1294   0.0604   1.0000
  15.250   1.7062   0.08935   0.08204  -0.1292   0.0567   1.0000
<< Back to GOE 225 (MVA H.35) AIRFOIL (goe225-il)

Polar data table (+)

Polar graphs


<< Back to GOE 225 (MVA H.35) AIRFOIL (goe225-il)