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FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il)
Reynolds number: 50,000
Max Cl/Cd: 4.98 at α=9.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66s171-il-50000.txt
Download as CSV file: xf-fx66s171-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-S-171 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.2457   0.11891   0.11329  -0.0381   1.0000   0.2950
 -10.000  -0.2483   0.11765   0.11227  -0.0350   1.0000   0.2990
  -9.750  -0.3114   0.12460   0.11956  -0.0245   1.0000   0.2968
  -9.500  -0.3118   0.12242   0.11739  -0.0307   0.9896   0.3087
  -9.250  -0.2636   0.11457   0.10944  -0.0359   0.9818   0.3145
  -9.000  -0.3451   0.09835   0.09335  -0.0578   0.9707   0.1725
  -8.750  -0.3729   0.08598   0.08101  -0.0691   0.9599   0.1411
  -8.500  -0.4316   0.07562   0.07054  -0.0803   0.9452   0.1266
  -8.250  -0.4398   0.07025   0.06502  -0.0840   0.9339   0.1234
  -8.000  -0.4629   0.06407   0.05842  -0.0880   0.9235   0.1184
  -7.750  -0.4820   0.05984   0.05332  -0.0892   0.9132   0.1142
  -7.500  -0.4782   0.05678   0.04982  -0.0890   0.9048   0.1144
  -7.250  -0.4567   0.05312   0.04604  -0.0904   0.8981   0.1173
  -7.000  -0.4431   0.05057   0.04317  -0.0903   0.8910   0.1189
  -6.750  -0.4271   0.04809   0.04029  -0.0901   0.8840   0.1209
  -6.500  -0.3977   0.04557   0.03718  -0.0917   0.8788   0.1268
  -6.250  -0.3954   0.04453   0.03586  -0.0889   0.8724   0.1303
  -6.000  -0.3679   0.04276   0.03403  -0.0893   0.8673   0.1380
  -5.750  -0.3416   0.04129   0.03244  -0.0894   0.8624   0.1500
  -5.500  -0.3331   0.04064   0.03160  -0.0868   0.8567   0.1598
  -5.250  -0.3097   0.03946   0.03062  -0.0863   0.8522   0.1792
  -5.000  -0.2825   0.03800   0.02939  -0.0866   0.8484   0.2175
  -4.750  -0.2851   0.03744   0.02931  -0.0830   0.8447   0.2548
  -4.500  -0.2844   0.03754   0.03087  -0.0782   0.8409   0.4072
  -4.250  -0.2789   0.04074   0.03415  -0.0711   0.8371   0.5100
  -4.000  -0.2641   0.04319   0.03653  -0.0654   0.8337   0.5580
  -3.750  -0.2660   0.04449   0.03777  -0.0599   0.8315   0.5858
  -3.500  -0.2671   0.04567   0.03891  -0.0542   0.8295   0.6109
  -3.250  -0.2664   0.04665   0.03984  -0.0488   0.8282   0.6357
  -3.000  -0.2642   0.04735   0.04044  -0.0445   0.8277   0.6619
  -2.750  -0.2640   0.04803   0.04109  -0.0391   0.8277   0.6842
  -2.500  -0.4348   0.04130   0.03484  -0.0247   1.0000   0.6010
  -2.250  -0.4284   0.04223   0.03571  -0.0199   1.0000   0.6333
  -2.000  -0.4218   0.04292   0.03634  -0.0155   1.0000   0.6638
  -1.750  -0.4155   0.04344   0.03679  -0.0110   1.0000   0.6938
  -1.500  -0.4094   0.04381   0.03709  -0.0067   1.0000   0.7234
  -1.250  -0.4051   0.04397   0.03721  -0.0020   1.0000   0.7493
  -1.000  -0.3968   0.04405   0.03718   0.0013   1.0000   0.7738
  -0.750  -0.3858   0.04413   0.03712   0.0035   1.0000   0.7947
  -0.500  -0.3706   0.04429   0.03710   0.0043   1.0000   0.8124
  -0.250  -0.3536   0.04450   0.03714   0.0045   1.0000   0.8273
   0.000  -0.3354   0.04479   0.03727   0.0042   1.0000   0.8415
   0.250  -0.3083   0.04570   0.03799   0.0021   0.9961   0.8559
   0.500  -0.2751   0.04701   0.03912  -0.0013   0.9877   0.8716
   0.750  -0.2416   0.04840   0.04037  -0.0048   0.9785   0.8894
   1.000  -0.2022   0.05018   0.04202  -0.0095   0.9684   0.9097
   1.250  -0.1510   0.05247   0.04421  -0.0168   0.9563   0.9341
   1.500  -0.1000   0.05424   0.04588  -0.0248   0.9423   0.9646
   1.750  -0.0630   0.05557   0.04704  -0.0303   0.9277   1.0000
   2.000  -0.0331   0.05678   0.04806  -0.0343   0.9141   1.0000
   2.250  -0.0026   0.05826   0.04936  -0.0383   0.9017   1.0000
   2.500   0.0346   0.06065   0.05153  -0.0435   0.8932   1.0000
   2.750   0.0702   0.06265   0.05335  -0.0480   0.8809   1.0000
   3.000   0.0955   0.06398   0.05451  -0.0505   0.8678   1.0000
   3.250   0.1209   0.06563   0.05600  -0.0530   0.8564   1.0000
   3.500   0.1630   0.06901   0.05918  -0.0581   0.8489   1.0000
   3.750   0.1800   0.06986   0.05991  -0.0588   0.8358   1.0000
   4.000   0.1985   0.07126   0.06118  -0.0596   0.8242   1.0000
   4.250   0.2292   0.07400   0.06378  -0.0624   0.8168   1.0000
   4.500   0.2500   0.07551   0.06519  -0.0633   0.8051   1.0000
   4.750   0.2639   0.07684   0.06644  -0.0634   0.7941   1.0000
   5.000   0.3024   0.08051   0.07000  -0.0669   0.7876   1.0000
   5.250   0.3086   0.08094   0.07039  -0.0657   0.7753   1.0000
   5.500   0.3225   0.08261   0.07200  -0.0658   0.7666   1.0000
   5.750   0.3520   0.08535   0.07468  -0.0678   0.7581   1.0000
   6.000   0.3594   0.08642   0.07573  -0.0670   0.7473   1.0000
   6.250   0.3945   0.09008   0.07936  -0.0697   0.7415   1.0000
   6.500   0.3956   0.09051   0.07978  -0.0682   0.7303   1.0000
   6.750   0.4197   0.09347   0.08272  -0.0696   0.7243   1.0000
   7.000   0.4321   0.09485   0.08409  -0.0694   0.7132   1.0000
   7.250   0.4430   0.09676   0.08603  -0.0693   0.7056   1.0000
   7.500   0.4662   0.09927   0.08855  -0.0704   0.6969   1.0000
   7.750   0.4723   0.10083   0.09013  -0.0698   0.6881   1.0000
   8.000   0.4994   0.10388   0.09320  -0.0714   0.6807   1.0000
   8.250   0.5016   0.10515   0.09450  -0.0705   0.6714   1.0000
   8.500   0.5317   0.10870   0.09811  -0.0723   0.6647   1.0000
   8.750   0.5303   0.10965   0.09909  -0.0712   0.6549   1.0000
   9.000   0.5617   0.11359   0.10310  -0.0732   0.6489   1.0000
   9.250   0.5581   0.11431   0.10387  -0.0720   0.6385   1.0000
   9.500   0.5918   0.11879   0.10844  -0.0742   0.6331   1.0000
   9.750   0.5848   0.11914   0.10884  -0.0729   0.6224   1.0000
  10.000   0.6157   0.12367   0.11345  -0.0748   0.6172   1.0000
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