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

GOE 746 AIRFOIL (goe746-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 746 AIRFOIL (goe746-il)
Reynolds number: 200,000
Max Cl/Cd: 73.84 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe746-il-200000-n5.txt
Download as CSV file: xf-goe746-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 746 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3873   0.09822   0.09401   0.0130   0.5946   0.0242
  -7.750  -0.3874   0.09503   0.09083   0.0106   0.5909   0.0243
  -7.500  -0.3873   0.09197   0.08775   0.0086   0.5876   0.0243
  -7.250  -0.3820   0.08837   0.08414   0.0057   0.5837   0.0244
  -7.000  -0.3297   0.07483   0.07079   0.0039   0.5716   0.0245
  -6.750  -0.3644   0.08060   0.07627   0.0000   0.5763   0.0245
  -6.250  -0.3433   0.07202   0.06757  -0.0042   0.5695   0.0250
  -6.000  -0.3320   0.06959   0.06513  -0.0032   0.5651   0.0258
  -5.750  -0.3174   0.06649   0.06196  -0.0045   0.5617   0.0267
  -5.500  -0.3009   0.06313   0.05849  -0.0064   0.5587   0.0276
  -5.250  -0.2717   0.05933   0.05441  -0.0110   0.5559   0.0315
  -5.000  -0.2478   0.05529   0.05012  -0.0130   0.5525   0.0319
  -4.750  -0.2306   0.05020   0.04484  -0.0140   0.5494   0.0323
  -4.500  -0.2160   0.04720   0.04180  -0.0140   0.5459   0.0332
  -4.250  -0.1966   0.04485   0.03932  -0.0140   0.5429   0.0341
  -4.000  -0.1745   0.04213   0.03644  -0.0143   0.5398   0.0351
  -3.500  -0.1245   0.03657   0.03032  -0.0145   0.5335   0.0401
  -3.250  -0.1025   0.03318   0.02666  -0.0140   0.5306   0.0363
  -3.000  -0.0778   0.03035   0.02344  -0.0134   0.5279   0.0360
  -2.750  -0.0526   0.02838   0.02117  -0.0128   0.5249   0.0375
  -2.500  -0.0269   0.02596   0.01838  -0.0121   0.5217   0.0364
  -2.250  -0.0006   0.02404   0.01611  -0.0114   0.5187   0.0360
  -2.000   0.0259   0.02264   0.01439  -0.0109   0.5158   0.0366
  -1.750   0.0528   0.02157   0.01303  -0.0104   0.5130   0.0380
  -1.500   0.0801   0.02040   0.01158  -0.0099   0.5102   0.0378
  -1.250   0.1079   0.01938   0.01033  -0.0096   0.5068   0.0377
  -1.000   0.1359   0.01853   0.00929  -0.0093   0.5036   0.0379
  -0.750   0.1638   0.01782   0.00841  -0.0091   0.5008   0.0381
  -0.500   0.1917   0.01722   0.00765  -0.0088   0.4981   0.0385
  -0.250   0.2194   0.01680   0.00710  -0.0086   0.4954   0.0397
   0.000   0.2474   0.01643   0.00670  -0.0085   0.4919   0.0407
   0.250   0.2751   0.01599   0.00624  -0.0083   0.4887   0.0409
   0.500   0.3023   0.01560   0.00582  -0.0081   0.4858   0.0411
   0.750   0.3293   0.01528   0.00547  -0.0078   0.4830   0.0413
   1.000   0.3560   0.01503   0.00519  -0.0075   0.4804   0.0416
   1.250   0.3827   0.01477   0.00498  -0.0072   0.4770   0.0423
   1.500   0.4094   0.01457   0.00479  -0.0070   0.4736   0.0433
   1.750   0.4362   0.01445   0.00464  -0.0067   0.4704   0.0449
   2.000   0.4632   0.01438   0.00453  -0.0065   0.4676   0.0467
   2.250   0.4902   0.01436   0.00448  -0.0063   0.4647   0.0511
   2.500   0.5174   0.01432   0.00452  -0.0061   0.4607   0.0569
   2.750   0.5443   0.01427   0.00452  -0.0059   0.4569   0.0737
   3.250   0.6821   0.01246   0.00473  -0.0233   0.4453   1.0000
   3.500   0.7079   0.01255   0.00479  -0.0230   0.4409   1.0000
   3.750   0.7335   0.01264   0.00483  -0.0225   0.4374   1.0000
   4.000   0.7592   0.01277   0.00495  -0.0222   0.4332   1.0000
   4.250   0.7850   0.01290   0.00511  -0.0219   0.4286   1.0000
   4.500   0.8105   0.01301   0.00522  -0.0215   0.4244   1.0000
   4.750   0.8360   0.01313   0.00530  -0.0211   0.4208   1.0000
   5.000   0.8616   0.01329   0.00554  -0.0208   0.4156   1.0000
   5.250   0.8870   0.01342   0.00572  -0.0205   0.4107   1.0000
   5.500   0.9122   0.01354   0.00581  -0.0201   0.4066   1.0000
   5.750   0.9376   0.01373   0.00611  -0.0198   0.4011   1.0000
   6.000   0.9628   0.01389   0.00631  -0.0195   0.3960   1.0000
   6.250   0.9878   0.01405   0.00649  -0.0191   0.3912   1.0000
   6.500   1.0129   0.01425   0.00681  -0.0188   0.3846   1.0000
   6.750   1.0377   0.01442   0.00701  -0.0184   0.3788   1.0000
   7.000   1.0624   0.01461   0.00730  -0.0181   0.3693   1.0000
   7.250   1.0866   0.01479   0.00751  -0.0177   0.3564   1.0000
   7.500   1.1105   0.01504   0.00777  -0.0173   0.3415   1.0000
   7.750   1.1332   0.01538   0.00806  -0.0169   0.3202   1.0000
   8.000   1.1552   0.01587   0.00852  -0.0165   0.2959   1.0000
   8.250   1.1752   0.01657   0.00912  -0.0160   0.2682   1.0000
   8.500   1.1932   0.01748   0.00992  -0.0155   0.2396   1.0000
   8.750   1.2061   0.01891   0.01112  -0.0148   0.2000   1.0000
   9.000   1.2122   0.02087   0.01278  -0.0139   0.1481   1.0000
   9.250   1.1960   0.02456   0.01595  -0.0124   0.0751   1.0000
   9.500   1.1764   0.02793   0.01917  -0.0107   0.0366   1.0000
   9.750   1.1671   0.03069   0.02193  -0.0095   0.0278   1.0000
  10.000   1.1663   0.03297   0.02429  -0.0088   0.0244   1.0000
  10.250   1.1668   0.03524   0.02667  -0.0083   0.0226   1.0000
  10.500   1.1665   0.03768   0.02921  -0.0079   0.0212   1.0000
  10.750   1.1639   0.04041   0.03204  -0.0077   0.0197   1.0000
  11.000   1.1591   0.04347   0.03522  -0.0075   0.0184   1.0000
  11.250   1.1588   0.04608   0.03794  -0.0075   0.0177   1.0000
  11.500   1.1577   0.04881   0.04080  -0.0074   0.0171   1.0000
  11.750   1.1548   0.05182   0.04393  -0.0076   0.0167   1.0000
  12.000   1.1512   0.05499   0.04722  -0.0078   0.0162   1.0000
  12.250   1.1476   0.05831   0.05068  -0.0082   0.0158   1.0000
  12.500   1.1427   0.06190   0.05439  -0.0087   0.0152   1.0000
  12.750   1.1385   0.06551   0.05811  -0.0094   0.0149   1.0000
  13.000   1.1338   0.06926   0.06196  -0.0101   0.0144   1.0000
  13.250   1.1298   0.07300   0.06580  -0.0109   0.0143   1.0000
  13.500   1.1241   0.07704   0.06993  -0.0118   0.0139   1.0000
  13.750   1.1185   0.08109   0.07405  -0.0127   0.0134   1.0000
  14.000   1.1109   0.08538   0.07839  -0.0137   0.0128   1.0000
  14.250   1.1078   0.08895   0.08199  -0.0143   0.0124   1.0000
  14.500   1.1102   0.09185   0.08496  -0.0148   0.0123   1.0000
<< Back to GOE 746 AIRFOIL (goe746-il)

Polar data table (+)

Polar graphs


<< Back to GOE 746 AIRFOIL (goe746-il)