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

GOE 427 AIRFOIL (goe427-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 427 AIRFOIL (goe427-il)
Reynolds number: 500,000
Max Cl/Cd: 114.45 at α=1.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe427-il-500000.txt
Download as CSV file: xf-goe427-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 427 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3547   0.09374   0.09160  -0.0178   1.0000   0.0120
  -7.500  -0.3630   0.09238   0.09029  -0.0153   1.0000   0.0120
  -7.250  -0.3726   0.09110   0.08906  -0.0127   1.0000   0.0124
  -7.000  -0.3598   0.08768   0.08565  -0.0161   0.9984   0.0129
  -6.750  -0.3187   0.08328   0.08121  -0.0291   0.9946   0.0145
  -6.500  -0.2868   0.07894   0.07685  -0.0383   0.9905   0.0147
  -6.250  -0.2543   0.07434   0.07222  -0.0460   0.9872   0.0147
  -6.000  -0.2199   0.06965   0.06747  -0.0537   0.9848   0.0148
  -2.750   0.2432   0.01568   0.01129  -0.1128   0.9408   0.0259
  -2.000   0.3293   0.01103   0.00563  -0.1122   0.9187   0.0325
  -1.750   0.3570   0.01013   0.00462  -0.1121   0.9085   0.0363
  -1.500   0.3854   0.00950   0.00387  -0.1119   0.8978   0.0388
  -1.250   0.4139   0.00908   0.00336  -0.1118   0.8873   0.0420
  -1.000   0.4410   0.00883   0.00303  -0.1114   0.8750   0.0438
  -0.750   0.4676   0.00830   0.00240  -0.1109   0.8623   0.0450
  -0.500   0.4940   0.00789   0.00194  -0.1104   0.8490   0.0474
  -0.250   0.5205   0.00769   0.00169  -0.1100   0.8353   0.0508
   0.000   0.5469   0.00756   0.00150  -0.1095   0.8208   0.0545
   0.250   0.5728   0.00749   0.00133  -0.1089   0.8029   0.0601
   0.500   0.5972   0.00741   0.00128  -0.1079   0.7760   0.1052
   0.750   0.6210   0.00729   0.00131  -0.1070   0.7484   0.2132
   1.000   0.6403   0.00644   0.00144  -0.1056   0.7246   0.6611
   1.250   0.6815   0.00597   0.00145  -0.1084   0.6967   1.0000
   1.500   0.7050   0.00616   0.00148  -0.1074   0.6654   1.0000
   1.750   0.7280   0.00638   0.00152  -0.1063   0.6262   1.0000
   2.000   0.7504   0.00667   0.00159  -0.1051   0.5849   1.0000
   2.250   0.7731   0.00698   0.00171  -0.1040   0.5524   1.0000
   2.500   0.7964   0.00728   0.00186  -0.1031   0.5241   1.0000
   2.750   0.8199   0.00758   0.00201  -0.1022   0.4958   1.0000
   3.000   0.8440   0.00782   0.00215  -0.1015   0.4692   1.0000
   3.250   0.8678   0.00809   0.00230  -0.1008   0.4363   1.0000
   3.500   0.8904   0.00847   0.00247  -0.0998   0.3868   1.0000
   3.750   0.9114   0.00904   0.00276  -0.0986   0.3248   1.0000
   4.000   0.9318   0.00972   0.00310  -0.0974   0.2538   1.0000
   4.250   0.9466   0.01109   0.00378  -0.0955   0.1162   1.0000
   4.500   0.9655   0.01205   0.00436  -0.0940   0.0469   1.0000
   4.750   0.9888   0.01248   0.00478  -0.0932   0.0412   1.0000
   5.000   1.0117   0.01297   0.00536  -0.0922   0.0367   1.0000
   5.250   1.0348   0.01341   0.00587  -0.0913   0.0353   1.0000
   5.500   1.0568   0.01395   0.00650  -0.0903   0.0329   1.0000
   5.750   1.0770   0.01469   0.00731  -0.0890   0.0293   1.0000
   6.000   1.0888   0.01638   0.00914  -0.0862   0.0254   1.0000
   6.250   1.1123   0.01670   0.00952  -0.0854   0.0238   1.0000
   6.500   1.1331   0.01732   0.01022  -0.0842   0.0212   1.0000
   6.750   1.1530   0.01800   0.01093  -0.0830   0.0189   1.0000
   7.000   1.1624   0.02048   0.01347  -0.0800   0.0168   1.0000
   7.250   1.1836   0.02116   0.01424  -0.0789   0.0159   1.0000
   7.500   1.2039   0.02218   0.01537  -0.0776   0.0149   1.0000
   7.750   1.2241   0.02318   0.01646  -0.0763   0.0136   1.0000
   8.000   1.2437   0.02373   0.01710  -0.0752   0.0124   1.0000
   8.250   1.2617   0.02480   0.01821  -0.0739   0.0115   1.0000
   8.500   1.2797   0.02996   0.02366  -0.0727   0.0108   1.0000
   8.750   1.2968   0.03134   0.02527  -0.0710   0.0105   1.0000
   9.000   1.3121   0.03380   0.02803  -0.0691   0.0102   1.0000
   9.250   1.3225   0.03738   0.03199  -0.0666   0.0100   1.0000
   9.500   1.3256   0.04166   0.03667  -0.0634   0.0102   1.0000
   9.750   1.3266   0.04768   0.04296  -0.0608   0.0110   1.0000
  10.000   1.3289   0.05021   0.04573  -0.0578   0.0111   1.0000
  10.250   1.3222   0.05376   0.04955  -0.0542   0.0111   1.0000
  10.500   1.3134   0.05628   0.05231  -0.0501   0.0111   1.0000
  10.750   1.2993   0.05900   0.05523  -0.0459   0.0112   1.0000
<< Back to GOE 427 AIRFOIL (goe427-il)

Polar data table (+)

Polar graphs


<< Back to GOE 427 AIRFOIL (goe427-il)