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

GOE 432 AIRFOIL (goe432-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 432 AIRFOIL (goe432-il)
Reynolds number: 100,000
Max Cl/Cd: 56.89 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe432-il-100000-n5.txt
Download as CSV file: xf-goe432-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 432 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3422   0.09090   0.08634  -0.0219   1.0000   0.0428
  -7.000  -0.3464   0.08868   0.08420  -0.0210   1.0000   0.0435
  -6.750  -0.3479   0.08614   0.08173  -0.0210   1.0000   0.0441
  -6.500  -0.3484   0.08347   0.07911  -0.0214   1.0000   0.0442
  -6.250  -0.3476   0.08061   0.07631  -0.0221   1.0000   0.0444
  -6.000  -0.3451   0.07760   0.07333  -0.0231   1.0000   0.0441
  -5.500  -0.3259   0.06600   0.06161  -0.0333   1.0000   0.0293
  -5.250  -0.3164   0.06281   0.05842  -0.0338   0.9997   0.0288
  -5.000  -0.2786   0.05707   0.05253  -0.0417   0.9937   0.0283
  -4.750  -0.2338   0.04979   0.04494  -0.0514   0.9875   0.0289
  -4.500  -0.1942   0.04325   0.03805  -0.0584   0.9814   0.0303
  -4.250  -0.1562   0.03835   0.03279  -0.0633   0.9765   0.0307
  -4.000  -0.1227   0.03437   0.02841  -0.0664   0.9700   0.0316
  -3.750  -0.0844   0.03116   0.02474  -0.0697   0.9652   0.0348
  -3.500  -0.0503   0.02742   0.02021  -0.0713   0.9582   0.0376
  -3.250  -0.0133   0.02479   0.01706  -0.0735   0.9529   0.0406
  -3.000   0.0189   0.02366   0.01569  -0.0745   0.9445   0.0460
  -2.750   0.0577   0.02180   0.01322  -0.0761   0.9383   0.0508
  -2.500   0.0892   0.02068   0.01203  -0.0769   0.9287   0.0573
  -2.250   0.1292   0.01963   0.01061  -0.0787   0.9221   0.0639
  -2.000   0.1588   0.01855   0.00952  -0.0789   0.9105   0.0688
  -1.750   0.1909   0.01773   0.00862  -0.0793   0.8995   0.0731
  -1.500   0.2242   0.01705   0.00787  -0.0800   0.8884   0.0801
  -1.250   0.2569   0.01639   0.00721  -0.0805   0.8763   0.0857
  -1.000   0.2895   0.01587   0.00659  -0.0810   0.8627   0.0918
  -0.750   0.3225   0.01534   0.00606  -0.0816   0.8475   0.1017
  -0.500   0.3564   0.01483   0.00558  -0.0823   0.8310   0.1239
  -0.250   0.3924   0.01405   0.00525  -0.0836   0.8125   0.2575
   0.000   0.4236   0.01277   0.00500  -0.0840   0.7905   0.6030
   0.500   0.5102   0.01213   0.00436  -0.0887   0.7274   1.0000
   0.750   0.5390   0.01226   0.00421  -0.0884   0.6933   1.0000
   1.000   0.5664   0.01244   0.00413  -0.0879   0.6608   1.0000
   1.250   0.5929   0.01266   0.00409  -0.0872   0.6305   1.0000
   1.500   0.6187   0.01290   0.00410  -0.0865   0.6037   1.0000
   1.750   0.6441   0.01317   0.00415  -0.0858   0.5805   1.0000
   2.000   0.6694   0.01345   0.00425  -0.0851   0.5606   1.0000
   2.250   0.6947   0.01374   0.00438  -0.0844   0.5435   1.0000
   2.500   0.7197   0.01404   0.00454  -0.0837   0.5267   1.0000
   2.750   0.7445   0.01435   0.00473  -0.0830   0.5102   1.0000
   3.000   0.7693   0.01465   0.00494  -0.0823   0.4953   1.0000
   3.250   0.7940   0.01494   0.00517  -0.0816   0.4813   1.0000
   3.500   0.8185   0.01523   0.00542  -0.0808   0.4667   1.0000
   3.750   0.8430   0.01552   0.00567  -0.0801   0.4530   1.0000
   4.000   0.8675   0.01581   0.00595  -0.0794   0.4406   1.0000
   4.250   0.8918   0.01610   0.00625  -0.0787   0.4280   1.0000
   4.500   0.9158   0.01641   0.00655  -0.0779   0.4147   1.0000
   4.750   0.9395   0.01672   0.00685  -0.0770   0.4004   1.0000
   5.000   0.9629   0.01704   0.00720  -0.0762   0.3850   1.0000
   5.250   0.9860   0.01737   0.00755  -0.0752   0.3689   1.0000
   5.500   1.0087   0.01773   0.00793  -0.0743   0.3516   1.0000
   5.750   1.0308   0.01813   0.00832  -0.0732   0.3328   1.0000
   6.000   1.0523   0.01858   0.00877  -0.0721   0.3131   1.0000
   6.250   1.0737   0.01905   0.00924  -0.0710   0.2921   1.0000
   6.500   1.0946   0.01956   0.00973  -0.0699   0.2713   1.0000
   6.750   1.1152   0.02012   0.01027  -0.0687   0.2540   1.0000
   7.250   1.1563   0.02131   0.01151  -0.0664   0.2278   1.0000
   7.500   1.1767   0.02194   0.01219  -0.0652   0.2177   1.0000
   7.750   1.1964   0.02263   0.01291  -0.0640   0.2094   1.0000
   8.000   1.2163   0.02332   0.01368  -0.0628   0.2022   1.0000
   8.250   1.2357   0.02408   0.01454  -0.0616   0.1962   1.0000
   8.500   1.2555   0.02484   0.01543  -0.0604   0.1900   1.0000
   8.750   1.2741   0.02574   0.01635  -0.0591   0.1847   1.0000
   9.000   1.2940   0.02656   0.01738  -0.0579   0.1788   1.0000
   9.250   1.3117   0.02749   0.01838  -0.0566   0.1716   1.0000
   9.500   1.3269   0.02833   0.01940  -0.0549   0.1612   1.0000
   9.750   1.3388   0.02911   0.02035  -0.0529   0.1486   1.0000
  10.000   1.3504   0.02975   0.02118  -0.0508   0.1345   1.0000
  10.250   1.3632   0.03031   0.02195  -0.0489   0.1179   1.0000
  10.500   1.3743   0.03103   0.02279  -0.0468   0.1005   1.0000
  10.750   1.3791   0.03226   0.02395  -0.0441   0.0815   1.0000
  11.000   1.3797   0.03403   0.02568  -0.0412   0.0572   1.0000
  11.250   1.3744   0.03635   0.02785  -0.0381   0.0410   1.0000
  11.500   1.3672   0.03893   0.03041  -0.0352   0.0339   1.0000
  11.750   1.3618   0.04149   0.03312  -0.0329   0.0302   1.0000
  12.000   1.3534   0.04448   0.03624  -0.0311   0.0280   1.0000
  12.250   1.3425   0.04795   0.03984  -0.0299   0.0267   1.0000
  12.500   1.3346   0.05142   0.04354  -0.0294   0.0254   1.0000
  12.750   1.3248   0.05546   0.04784  -0.0296   0.0242   1.0000
  13.000   1.3137   0.05998   0.05257  -0.0305   0.0234   1.0000
  13.250   1.3015   0.06506   0.05784  -0.0321   0.0226   1.0000
  13.500   1.2883   0.07063   0.06360  -0.0342   0.0222   1.0000
  13.750   1.2739   0.07666   0.06982  -0.0367   0.0217   1.0000
  14.000   1.2589   0.08298   0.07631  -0.0395   0.0215   1.0000
  14.250   1.2437   0.08958   0.08307  -0.0425   0.0213   1.0000
<< Back to GOE 432 AIRFOIL (goe432-il)

Polar data table (+)

Polar graphs


<< Back to GOE 432 AIRFOIL (goe432-il)