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

GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)
Reynolds number: 1,000,000
Max Cl/Cd: 86.7 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe417a-il-1000000.txt
Download as CSV file: xf-goe417a-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 417A (GEW. PLATTE) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2270   0.10037   0.09874  -0.0390   0.9921   0.0129
  -9.500  -0.2178   0.09702   0.09539  -0.0411   0.9907   0.0129
  -9.250  -0.2088   0.09372   0.09210  -0.0435   0.9894   0.0130
  -9.000  -0.1995   0.08932   0.08770  -0.0451   0.9882   0.0131
  -8.750  -0.1887   0.08578   0.08415  -0.0469   0.9872   0.0131
  -8.500  -0.1771   0.08256   0.08093  -0.0489   0.9863   0.0133
  -8.250  -0.2617   0.09513   0.09343  -0.0420   0.9879   0.0131
  -8.000  -0.2459   0.09204   0.09033  -0.0451   0.9863   0.0133
  -7.750  -0.2290   0.08916   0.08745  -0.0487   0.9849   0.0135
  -7.500  -0.2103   0.08620   0.08449  -0.0527   0.9837   0.0138
  -7.250  -0.1947   0.08338   0.08166  -0.0555   0.9806   0.0142
  -7.000  -0.1781   0.08057   0.07885  -0.0585   0.9774   0.0147
  -6.750  -0.1562   0.07748   0.07575  -0.0628   0.9753   0.0152
  -6.500  -0.1302   0.07425   0.07250  -0.0680   0.9737   0.0155
  -6.250  -0.0996   0.07098   0.06920  -0.0745   0.9724   0.0156
  -6.000  -0.0660   0.06760   0.06579  -0.0817   0.9714   0.0156
  -5.750  -0.0294   0.06420   0.06234  -0.0895   0.9706   0.0157
  -5.250   0.0117   0.05795   0.05606  -0.0947   0.9635   0.0158
  -5.000   0.0371   0.05489   0.05298  -0.0979   0.9618   0.0159
  -4.750   0.0678   0.05210   0.05016  -0.1023   0.9604   0.0161
  -4.500   0.1004   0.04948   0.04752  -0.1068   0.9589   0.0164
  -4.250   0.1343   0.04685   0.04485  -0.1114   0.9572   0.0168
  -4.000   0.1671   0.04436   0.04231  -0.1155   0.9556   0.0175
  -3.750   0.1876   0.04246   0.04038  -0.1160   0.9489   0.0182
  -3.500   0.2246   0.04034   0.03817  -0.1198   0.9450   0.0185
  -3.250   0.2600   0.03816   0.03591  -0.1233   0.9416   0.0186
  -3.000   0.2848   0.03638   0.03406  -0.1240   0.9361   0.0186
  -2.750   0.3133   0.03450   0.03209  -0.1252   0.9306   0.0186
  -2.000   0.3850   0.02821   0.02560  -0.1262   0.9124   0.0188
  -1.750   0.4049   0.02663   0.02399  -0.1258   0.9056   0.0191
  -1.500   0.4286   0.02525   0.02255  -0.1256   0.8979   0.0193
  -1.250   0.4522   0.02397   0.02121  -0.1252   0.8887   0.0198
  -1.000   0.4775   0.02268   0.01982  -0.1249   0.8778   0.0206
  -0.750   0.5042   0.02161   0.01862  -0.1242   0.8652   0.0216
  -0.500   0.5314   0.02069   0.01755  -0.1233   0.8481   0.0218
  -0.250   0.5551   0.01977   0.01642  -0.1219   0.8159   0.0219
   0.000   0.5732   0.01903   0.01536  -0.1195   0.7641   0.0219
   0.250   0.5919   0.01828   0.01437  -0.1173   0.7300   0.0219
   0.500   0.6122   0.01748   0.01338  -0.1155   0.7054   0.0219
   0.750   0.6302   0.01597   0.01166  -0.1134   0.6824   0.0222
   1.000   0.6495   0.01523   0.01080  -0.1119   0.6586   0.0224
   1.250   0.6695   0.01466   0.01008  -0.1103   0.6323   0.0227
   1.500   0.6893   0.01419   0.00942  -0.1085   0.6006   0.0232
   1.750   0.7100   0.01373   0.00878  -0.1067   0.5727   0.0241
   2.000   0.7347   0.01370   0.00850  -0.1048   0.5497   0.0258
   2.250   0.7557   0.01332   0.00791  -0.1030   0.5234   0.0258
   5.750   1.0418   0.01289   0.00478  -0.0813   0.0408   0.0346
   6.000   1.0614   0.01306   0.00493  -0.0795   0.0343   0.0332
   6.250   1.0821   0.01324   0.00511  -0.0780   0.0315   0.0335
   6.500   1.1010   0.01358   0.00547  -0.0762   0.0277   0.0346
   6.750   1.1214   0.01383   0.00575  -0.0746   0.0261   0.0365
   7.000   1.1415   0.01407   0.00601  -0.0731   0.0246   0.0395
   7.250   1.1610   0.01437   0.00631  -0.0714   0.0229   0.0431
   7.500   1.1766   0.01488   0.00699  -0.0690   0.0208   0.1548
   7.750   1.3638   0.01573   0.00908  -0.1068   0.0170   1.0000
   8.000   1.3771   0.01657   0.00999  -0.1040   0.0160   1.0000
   8.250   1.3957   0.01694   0.01038  -0.1022   0.0156   1.0000
   8.500   1.4131   0.01738   0.01086  -0.1003   0.0151   1.0000
   8.750   1.4297   0.01786   0.01137  -0.0982   0.0145   1.0000
   9.000   1.4457   0.01836   0.01192  -0.0960   0.0140   1.0000
   9.250   1.4610   0.01887   0.01246  -0.0937   0.0135   1.0000
   9.500   1.4746   0.01948   0.01310  -0.0911   0.0131   1.0000
   9.750   1.4808   0.02056   0.01426  -0.0871   0.0126   1.0000
  10.000   1.4797   0.02180   0.01560  -0.0818   0.0123   1.0000
  10.250   1.4893   0.02229   0.01616  -0.0784   0.0121   1.0000
  10.500   1.4998   0.02276   0.01668  -0.0753   0.0118   1.0000
  10.750   1.5073   0.02351   0.01752  -0.0718   0.0116   1.0000
  11.000   1.5175   0.02409   0.01815  -0.0688   0.0112   1.0000
  11.250   1.5273   0.02473   0.01886  -0.0659   0.0109   1.0000
  11.500   1.5340   0.02565   0.01986  -0.0626   0.0106   1.0000
  11.750   1.5439   0.02630   0.02055  -0.0599   0.0103   1.0000
  12.000   1.5492   0.02738   0.02172  -0.0567   0.0102   1.0000
  12.250   1.5554   0.02838   0.02279  -0.0537   0.0100   1.0000
  12.500   1.5602   0.02954   0.02402  -0.0507   0.0099   1.0000
  12.750   1.5615   0.03104   0.02560  -0.0474   0.0096   1.0000
  13.000   1.5561   0.03338   0.02809  -0.0436   0.0094   1.0000
  13.250   1.5490   0.03613   0.03103  -0.0399   0.0092   1.0000
  13.500   1.5526   0.03753   0.03254  -0.0375   0.0091   1.0000
  13.750   1.5554   0.03905   0.03418  -0.0353   0.0090   1.0000
  14.000   1.5527   0.04137   0.03664  -0.0328   0.0089   1.0000
  14.250   1.5506   0.04362   0.03903  -0.0307   0.0088   1.0000
  14.500   1.5476   0.04608   0.04163  -0.0289   0.0087   1.0000
  14.750   1.5439   0.04871   0.04440  -0.0275   0.0086   1.0000
  15.000   1.5367   0.05195   0.04779  -0.0262   0.0086   1.0000
  15.250   1.5275   0.05562   0.05162  -0.0254   0.0085   1.0000
  15.500   1.5210   0.05905   0.05518  -0.0251   0.0083   1.0000
  15.750   1.5065   0.06384   0.06015  -0.0252   0.0084   1.0000
  16.000   1.4902   0.06920   0.06571  -0.0259   0.0083   1.0000
  16.250   1.4798   0.07390   0.07054  -0.0270   0.0082   1.0000
  16.500   1.4693   0.07888   0.07564  -0.0285   0.0081   1.0000
  16.750   1.4491   0.08574   0.08268  -0.0309   0.0081   1.0000
  17.000   1.4323   0.09227   0.08936  -0.0336   0.0080   1.0000
  17.250   1.4026   0.10161   0.09891  -0.0377   0.0081   1.0000
  17.500   1.3868   0.10873   0.10616  -0.0413   0.0081   1.0000
  17.750   1.3653   0.11735   0.11493  -0.0458   0.0081   1.0000
  18.000   1.3271   0.13057   0.12837  -0.0533   0.0082   1.0000
  18.250   1.2825   0.14663   0.14466  -0.0628   0.0084   1.0000
<< Back to GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)

Polar data table (+)

Polar graphs


<< Back to GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)