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

GOE 571 AIRFOIL (goe571-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 571 AIRFOIL (goe571-il)
Reynolds number: 1,000,000
Max Cl/Cd: 52.53 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe571-il-1000000-n5.txt
Download as CSV file: xf-goe571-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 571 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.000   0.5113   0.09494   0.09023  -0.1697   0.4907   0.0185
 -11.750   0.5173   0.09259   0.08787  -0.1705   0.4862   0.0186
 -11.500   0.5229   0.09025   0.08550  -0.1713   0.4818   0.0185
 -11.250   0.5286   0.08795   0.08319  -0.1720   0.4777   0.0185
 -11.000   0.5303   0.08501   0.08025  -0.1729   0.4749   0.0189
 -10.750   0.5310   0.08213   0.07738  -0.1737   0.4715   0.0197
 -10.250   0.5296   0.07599   0.07123  -0.1752   0.4645   0.0191
 -10.000   0.5188   0.07157   0.06681  -0.1763   0.4624   0.0194
  -9.750   0.5377   0.07192   0.06716  -0.1761   0.4574   0.0197
  -9.500   0.5169   0.06635   0.06162  -0.1772   0.4556   0.0195
  -9.250   0.5117   0.06342   0.05870  -0.1776   0.4535   0.0195
  -9.000   0.1527   0.03097   0.02601  -0.1661   0.4705   0.0195
  -8.750   0.1796   0.03222   0.02735  -0.1650   0.4650   0.0196
  -8.500   0.1292   0.02937   0.02428  -0.1539   0.4630   0.0196
  -8.250   0.1292   0.02894   0.02380  -0.1494   0.4591   0.0197
  -8.000   0.1053   0.02708   0.02176  -0.1414   0.4571   0.0197
  -7.750   0.0977   0.02596   0.02051  -0.1356   0.4541   0.0198
  -7.500   0.1061   0.02571   0.02022  -0.1323   0.4502   0.0199
  -7.250   0.1040   0.02473   0.01912  -0.1273   0.4467   0.0199
  -6.750   0.1102   0.02339   0.01757  -0.1187   0.4392   0.0201
  -6.500   0.1149   0.02264   0.01671  -0.1146   0.4365   0.0201
  -6.250   0.1220   0.02201   0.01598  -0.1109   0.4335   0.0202
  -6.000   0.1323   0.02164   0.01554  -0.1077   0.4303   0.0203
  -5.750   0.1391   0.02106   0.01484  -0.1039   0.4273   0.0204
  -5.250   0.1527   0.01966   0.01315  -0.0961   0.4225   0.0205
  -5.000   0.1672   0.01943   0.01288  -0.0936   0.4205   0.0206
  -4.750   0.1794   0.01897   0.01231  -0.0906   0.4185   0.0207
  -4.500   0.1916   0.01860   0.01185  -0.0877   0.4161   0.0208
  -4.250   0.2035   0.01822   0.01137  -0.0846   0.4138   0.0209
  -4.000   0.2166   0.01795   0.01101  -0.0818   0.4113   0.0211
  -3.750   0.2295   0.01758   0.01055  -0.0790   0.4090   0.0210
  -3.500   0.2425   0.01736   0.01025  -0.0762   0.4063   0.0212
  -3.250   0.2580   0.01716   0.01000  -0.0738   0.4046   0.0213
  -3.000   0.2739   0.01688   0.00964  -0.0715   0.4031   0.0214
  -2.750   0.2901   0.01661   0.00932  -0.0693   0.4014   0.0214
  -2.500   0.3064   0.01644   0.00911  -0.0672   0.3998   0.0215
  -2.250   0.3219   0.01635   0.00896  -0.0649   0.3979   0.0216
  -2.000   0.3373   0.01618   0.00874  -0.0626   0.3960   0.0217
  -1.750   0.3530   0.01606   0.00858  -0.0604   0.3942   0.0219
  -1.500   0.3681   0.01606   0.00854  -0.0582   0.3923   0.0220
  -1.250   0.3825   0.01612   0.00856  -0.0558   0.3902   0.0222
  -1.000   0.3977   0.01603   0.00844  -0.0536   0.3883   0.0221
  -0.750   0.4160   0.01606   0.00844  -0.0520   0.3874   0.0224
  -0.500   0.4336   0.01605   0.00841  -0.0503   0.3863   0.0225
  -0.250   0.4514   0.01600   0.00836  -0.0486   0.3852   0.0224
   0.000   0.4694   0.01602   0.00837  -0.0471   0.3834   0.0225
   0.250   0.4870   0.01593   0.00829  -0.0455   0.3820   0.0229
   0.500   0.5041   0.01601   0.00836  -0.0438   0.3805   0.0230
   0.750   0.5217   0.01612   0.00847  -0.0424   0.3789   0.0233
   1.000   0.5400   0.01624   0.00858  -0.0410   0.3773   0.0233
   1.250   0.5564   0.01644   0.00878  -0.0395   0.3755   0.0236
   1.500   0.5740   0.01664   0.00897  -0.0382   0.3741   0.0237
   1.750   0.5910   0.01690   0.00922  -0.0369   0.3717   0.0239
   2.000   0.6121   0.01706   0.00939  -0.0363   0.3708   0.0241
   2.250   0.6340   0.01724   0.00957  -0.0358   0.3697   0.0244
   2.500   0.6560   0.01744   0.00979  -0.0355   0.3687   0.0246
   2.750   0.6784   0.01766   0.01001  -0.0352   0.3675   0.0250
   3.000   0.7011   0.01790   0.01027  -0.0351   0.3664   0.0252
   3.250   0.7233   0.01820   0.01057  -0.0350   0.3651   0.0255
   3.500   0.7462   0.01850   0.01088  -0.0350   0.3639   0.0256
   3.750   0.7703   0.01882   0.01120  -0.0354   0.3621   0.0264
   4.000   0.7913   0.01928   0.01167  -0.0354   0.3602   0.0263
   4.250   0.8136   0.01973   0.01213  -0.0357   0.3584   0.0269
   4.500   0.8359   0.02022   0.01262  -0.0361   0.3568   0.0277
   4.750   0.8562   0.02081   0.01322  -0.0362   0.3550   0.0277
   5.000   0.8837   0.02118   0.01361  -0.0374   0.3544   0.0290
   5.250   0.9112   0.02157   0.01403  -0.0386   0.3532   0.0288
   5.500   0.9389   0.02200   0.01449  -0.0400   0.3522   0.0310
   5.750   0.9623   0.02253   0.01504  -0.0407   0.3509   0.0320
   6.250   1.0071   0.02361   0.01616  -0.0417   0.3481   0.0371
   6.500   1.0304   0.02426   0.01695  -0.0428   0.3464   0.0989
   7.000   1.3163   0.02506   0.01979  -0.0932   0.3413   1.0000
   7.250   1.3259   0.02610   0.02083  -0.0918   0.3394   1.0000
   7.500   1.3413   0.02682   0.02157  -0.0912   0.3384   1.0000
   7.750   1.3554   0.02764   0.02240  -0.0905   0.3376   1.0000
   8.000   1.3701   0.02842   0.02320  -0.0898   0.3363   1.0000
   8.250   1.3822   0.02941   0.02420  -0.0890   0.3343   1.0000
   8.500   1.3942   0.03040   0.02520  -0.0882   0.3326   1.0000
   8.750   1.4053   0.03147   0.02628  -0.0873   0.3311   1.0000
   9.000   1.4127   0.03283   0.02764  -0.0863   0.3286   1.0000
   9.250   1.4230   0.03399   0.02881  -0.0854   0.3276   1.0000
   9.500   1.4278   0.03558   0.03040  -0.0843   0.3253   1.0000
   9.750   1.4381   0.03676   0.03160  -0.0835   0.3238   1.0000
  10.000   1.4493   0.03791   0.03278  -0.0829   0.3227   1.0000
  10.250   1.4608   0.03903   0.03393  -0.0824   0.3213   1.0000
  10.500   1.4720   0.04019   0.03511  -0.0819   0.3195   1.0000
  10.750   1.4801   0.04161   0.03655  -0.0812   0.3178   1.0000
  11.000   1.4865   0.04317   0.03813  -0.0804   0.3154   1.0000
  11.250   1.4933   0.04470   0.03967  -0.0796   0.3134   1.0000
  11.500   1.4940   0.04676   0.04174  -0.0785   0.3112   1.0000
  11.750   1.5009   0.04830   0.04330  -0.0779   0.3092   1.0000
  12.000   1.5106   0.04963   0.04467  -0.0775   0.3079   1.0000
  12.250   1.5179   0.05120   0.04628  -0.0769   0.3060   1.0000
  12.500   1.5257   0.05270   0.04780  -0.0764   0.3039   1.0000
  12.750   1.5299   0.05456   0.04967  -0.0758   0.3009   1.0000
  13.000   1.5274   0.05700   0.05214  -0.0748   0.2980   1.0000
  13.250   1.5280   0.05918   0.05432  -0.0740   0.2954   1.0000
  13.500   1.5309   0.06120   0.05638  -0.0734   0.2924   1.0000
  13.750   1.5343   0.06318   0.05839  -0.0729   0.2891   1.0000
  14.000   1.5391   0.06502   0.06026  -0.0724   0.2872   1.0000
  14.250   1.5250   0.06872   0.06397  -0.0712   0.2823   1.0000
  14.500   1.5281   0.07075   0.06603  -0.0707   0.2792   1.0000
  14.750   1.5234   0.07359   0.06891  -0.0700   0.2752   1.0000
  15.000   1.5228   0.07603   0.07137  -0.0695   0.2724   1.0000
  15.250   1.5060   0.08016   0.07552  -0.0684   0.2678   1.0000
  15.500   1.5083   0.08236   0.07775  -0.0681   0.2643   1.0000
  15.750   1.5056   0.08509   0.08052  -0.0676   0.2614   1.0000
  16.000   1.4979   0.08835   0.08380  -0.0671   0.2577   1.0000
  16.250   1.4843   0.09229   0.08777  -0.0664   0.2536   1.0000
  16.500   1.4816   0.09508   0.09060  -0.0661   0.2496   1.0000
  16.750   1.4734   0.09851   0.09404  -0.0657   0.2453   1.0000
  17.000   1.4638   0.10208   0.09764  -0.0653   0.2425   1.0000
  17.250   1.4602   0.10500   0.10059  -0.0652   0.2383   1.0000
<< Back to GOE 571 AIRFOIL (goe571-il)

Polar data table (+)

Polar graphs


<< Back to GOE 571 AIRFOIL (goe571-il)