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GOE 101 AIRFOIL (goe101-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 101 AIRFOIL (goe101-il)
Reynolds number: 50,000
Max Cl/Cd: 41.13 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe101-il-50000-n5.txt
Download as CSV file: xf-goe101-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 101 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3594   0.09896   0.09241  -0.0193   1.0000   0.0792
  -7.500  -0.3609   0.09699   0.09055  -0.0209   1.0000   0.0821
  -7.250  -0.3625   0.09576   0.08944  -0.0267   1.0000   0.0840
  -7.000  -0.3573   0.09269   0.08647  -0.0302   1.0000   0.0849
  -6.750  -0.3470   0.08748   0.08125  -0.0247   1.0000   0.0889
  -6.500  -0.3406   0.08450   0.07833  -0.0257   1.0000   0.0931
  -6.250  -0.3336   0.08270   0.07656  -0.0331   1.0000   0.0984
  -5.750  -0.3197   0.07539   0.06940  -0.0306   1.0000   0.1089
  -5.500  -0.3102   0.07228   0.06631  -0.0343   1.0000   0.1160
  -5.250  -0.2996   0.06966   0.06368  -0.0356   1.0000   0.1271
  -5.000  -0.2920   0.06610   0.06020  -0.0348   1.0000   0.1330
  -4.750  -0.2784   0.06328   0.05732  -0.0378   1.0000   0.1450
  -4.500  -0.2676   0.06046   0.05449  -0.0375   1.0000   0.1562
  -4.250  -0.2582   0.05733   0.05141  -0.0367   1.0000   0.1651
  -4.000  -0.2028   0.05138   0.04450  -0.0471   1.0000   0.0843
  -3.750  -0.1920   0.04785   0.04108  -0.0464   1.0000   0.0867
  -3.500  -0.1731   0.04474   0.03777  -0.0466   1.0000   0.0783
  -3.250  -0.1252   0.04020   0.03255  -0.0520   0.9913   0.0679
  -3.000  -0.0799   0.03652   0.02833  -0.0567   0.9803   0.0673
  -2.500   0.0094   0.03109   0.02149  -0.0639   0.9564   0.0701
  -2.250   0.0506   0.02874   0.01865  -0.0665   0.9438   0.0709
  -2.000   0.0899   0.02700   0.01676  -0.0693   0.9306   0.0785
  -1.750   0.1312   0.02544   0.01477  -0.0714   0.9172   0.0795
  -1.500   0.1705   0.02415   0.01310  -0.0731   0.9028   0.0800
  -1.250   0.2080   0.02310   0.01176  -0.0742   0.8876   0.0809
  -1.000   0.2457   0.02223   0.01061  -0.0754   0.8720   0.0823
  -0.750   0.2814   0.02154   0.00964  -0.0762   0.8558   0.0845
  -0.500   0.3141   0.02087   0.00889  -0.0768   0.8383   0.0882
  -0.250   0.3453   0.02042   0.00829  -0.0770   0.8200   0.0951
   0.000   0.3762   0.02001   0.00783  -0.0772   0.8025   0.1140
   0.250   0.4064   0.01939   0.00739  -0.0774   0.7858   0.1569
   0.500   0.4396   0.01724   0.00704  -0.0777   0.7704   1.0000
   0.750   0.4676   0.01743   0.00685  -0.0773   0.7543   1.0000
   1.000   0.4950   0.01765   0.00678  -0.0768   0.7390   1.0000
   1.250   0.5217   0.01791   0.00682  -0.0763   0.7241   1.0000
   1.500   0.5482   0.01819   0.00692  -0.0758   0.7101   1.0000
   1.750   0.5742   0.01851   0.00710  -0.0753   0.6966   1.0000
   2.000   0.6001   0.01885   0.00733  -0.0749   0.6838   1.0000
   2.250   0.6263   0.01921   0.00761  -0.0745   0.6717   1.0000
   2.500   0.6528   0.01957   0.00794  -0.0742   0.6607   1.0000
   2.750   0.6798   0.01992   0.00823  -0.0739   0.6509   1.0000
   3.000   0.7054   0.02036   0.00868  -0.0736   0.6398   1.0000
   3.250   0.7313   0.02080   0.00914  -0.0733   0.6297   1.0000
   3.500   0.7582   0.02118   0.00956  -0.0730   0.6212   1.0000
   3.750   0.7830   0.02168   0.01016  -0.0726   0.6106   1.0000
   4.000   0.8084   0.02216   0.01073  -0.0722   0.6011   1.0000
   4.250   0.8346   0.02254   0.01120  -0.0717   0.5917   1.0000
   4.500   0.8588   0.02294   0.01170  -0.0709   0.5784   1.0000
   4.750   0.8828   0.02327   0.01212  -0.0699   0.5640   1.0000
   5.000   0.9067   0.02362   0.01258  -0.0690   0.5496   1.0000
   5.250   0.9305   0.02404   0.01321  -0.0681   0.5362   1.0000
   5.500   0.9543   0.02452   0.01389  -0.0673   0.5237   1.0000
   5.750   0.9781   0.02498   0.01456  -0.0665   0.5105   1.0000
   6.000   1.0014   0.02540   0.01518  -0.0654   0.4957   1.0000
   6.250   1.0243   0.02577   0.01577  -0.0643   0.4793   1.0000
   6.500   1.0471   0.02607   0.01633  -0.0630   0.4613   1.0000
   6.750   1.0681   0.02645   0.01699  -0.0616   0.4407   1.0000
   7.000   1.0885   0.02668   0.01744  -0.0598   0.4164   1.0000
   7.250   1.1032   0.02682   0.01776  -0.0574   0.3754   1.0000
   7.500   1.1158   0.02714   0.01814  -0.0550   0.3168   1.0000
   7.750   1.1225   0.02826   0.01873  -0.0521   0.2302   1.0000
   8.000   1.1241   0.03071   0.02055  -0.0497   0.1485   1.0000
   8.250   1.1231   0.03366   0.02298  -0.0474   0.0864   1.0000
   8.500   1.1222   0.03649   0.02566  -0.0450   0.0626   1.0000
   8.750   1.1239   0.03888   0.02808  -0.0428   0.0512   1.0000
   9.000   1.1227   0.04125   0.03052  -0.0404   0.0464   1.0000
   9.250   1.1222   0.04367   0.03308  -0.0384   0.0435   1.0000
   9.500   1.1205   0.04635   0.03587  -0.0369   0.0410   1.0000
   9.750   1.1184   0.04928   0.03887  -0.0357   0.0388   1.0000
  10.000   1.1211   0.05192   0.04175  -0.0346   0.0367   1.0000
  10.250   1.1235   0.05472   0.04472  -0.0338   0.0343   1.0000
  10.500   1.1264   0.05757   0.04768  -0.0330   0.0325   1.0000
  10.750   1.1333   0.06029   0.05051  -0.0317   0.0313   1.0000
  11.000   1.1491   0.06270   0.05319  -0.0297   0.0304   1.0000
  11.250   1.1605   0.06578   0.05675  -0.0285   0.0297   1.0000
  11.500   1.1636   0.06954   0.06088  -0.0281   0.0290   1.0000
  11.750   1.1607   0.07382   0.06550  -0.0284   0.0284   1.0000
  12.000   1.1537   0.07859   0.07059  -0.0294   0.0281   1.0000
  12.250   1.1435   0.08387   0.07615  -0.0311   0.0278   1.0000
  12.500   1.1312   0.08967   0.08221  -0.0335   0.0278   1.0000
  12.750   1.1170   0.09605   0.08883  -0.0366   0.0279   1.0000
  13.000   1.1012   0.10306   0.09604  -0.0404   0.0280   1.0000
  13.250   1.0855   0.11049   0.10365  -0.0446   0.0284   1.0000
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