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

FX 66-S-171 AIRFOIL (fx66s171-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: FX 66-S-171 AIRFOIL (fx66s171-il)
Reynolds number: 500,000
Max Cl/Cd: 113.93 at α=9°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66s171-il-500000.txt
Download as CSV file: xf-fx66s171-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.4244   0.09427   0.09209  -0.0534   1.0000   0.0151
 -12.750  -0.4702   0.07971   0.07745  -0.0618   1.0000   0.0146
 -12.500  -0.5068   0.06903   0.06662  -0.0692   1.0000   0.0144
 -12.250  -0.5374   0.06098   0.05845  -0.0746   1.0000   0.0141
 -12.000  -0.5457   0.04825   0.04525  -0.0935   0.9330   0.0137
 -11.750  -0.5268   0.03867   0.03483  -0.1125   0.8532   0.0137
 -11.500  -0.5412   0.03551   0.03120  -0.1122   0.8009   0.0136
 -11.250  -0.5502   0.03320   0.02857  -0.1112   0.7708   0.0137
 -11.000  -0.5581   0.03110   0.02619  -0.1094   0.7495   0.0137
 -10.750  -0.5601   0.02991   0.02478  -0.1069   0.7325   0.0140
 -10.500  -0.5581   0.02835   0.02295  -0.1047   0.7190   0.0142
 -10.250  -0.5497   0.02699   0.02135  -0.1031   0.7073   0.0145
 -10.000  -0.5379   0.02573   0.01985  -0.1017   0.6966   0.0148
  -9.750  -0.5232   0.02461   0.01852  -0.1005   0.6870   0.0151
  -9.500  -0.5066   0.02363   0.01732  -0.0995   0.6785   0.0152
  -9.250  -0.4928   0.02148   0.01503  -0.0982   0.6707   0.0157
  -9.000  -0.4752   0.02048   0.01392  -0.0973   0.6635   0.0161
  -8.750  -0.4555   0.01969   0.01308  -0.0965   0.6562   0.0167
  -8.500  -0.4352   0.01902   0.01230  -0.0957   0.6497   0.0173
  -8.250  -0.4142   0.01833   0.01152  -0.0950   0.6435   0.0180
  -8.000  -0.3925   0.01772   0.01080  -0.0943   0.6374   0.0189
  -7.750  -0.3698   0.01725   0.01020  -0.0938   0.6320   0.0195
  -7.500  -0.3519   0.01610   0.00902  -0.0928   0.6267   0.0209
  -7.250  -0.3280   0.01571   0.00858  -0.0924   0.6215   0.0228
  -7.000  -0.3035   0.01537   0.00813  -0.0921   0.6168   0.0245
  -6.750  -0.2818   0.01456   0.00730  -0.0915   0.6121   0.0272
  -6.500  -0.2566   0.01421   0.00689  -0.0913   0.6074   0.0300
  -6.250  -0.2327   0.01370   0.00628  -0.0909   0.6032   0.0334
  -6.000  -0.2069   0.01336   0.00591  -0.0908   0.5992   0.0376
  -5.750  -0.1815   0.01291   0.00544  -0.0906   0.5953   0.0436
  -5.500  -0.1556   0.01255   0.00506  -0.0904   0.5916   0.0530
  -5.250  -0.1300   0.01215   0.00468  -0.0903   0.5883   0.0720
  -5.000  -0.1052   0.01156   0.00429  -0.0902   0.5851   0.1195
  -4.750  -0.0811   0.01079   0.00387  -0.0902   0.5818   0.2064
  -4.500  -0.0571   0.01005   0.00355  -0.0901   0.5784   0.3230
  -4.250  -0.0304   0.00980   0.00343  -0.0901   0.5751   0.3766
  -4.000  -0.0027   0.00978   0.00342  -0.0902   0.5720   0.4096
  -3.750   0.0258   0.00978   0.00341  -0.0903   0.5691   0.4323
  -3.500   0.0542   0.00980   0.00341  -0.0904   0.5661   0.4499
  -3.250   0.0828   0.00985   0.00343  -0.0905   0.5633   0.4644
  -3.000   0.1116   0.00993   0.00343  -0.0906   0.5607   0.4763
  -2.750   0.1400   0.01001   0.00345  -0.0907   0.5580   0.4857
  -2.500   0.1685   0.01012   0.00351  -0.0909   0.5554   0.4968
  -2.250   0.1972   0.01019   0.00354  -0.0910   0.5530   0.5068
  -2.000   0.2255   0.01020   0.00356  -0.0911   0.5506   0.5142
  -1.500   0.2827   0.01027   0.00358  -0.0914   0.5460   0.5278
  -1.250   0.3114   0.01032   0.00357  -0.0916   0.5439   0.5325
  -1.000   0.3404   0.01042   0.00358  -0.0919   0.5417   0.5368
  -0.750   0.3689   0.01042   0.00359  -0.0921   0.5397   0.5406
  -0.500   0.3973   0.01041   0.00360  -0.0923   0.5373   0.5444
  -0.250   0.4258   0.01043   0.00360  -0.0924   0.5347   0.5486
   0.000   0.4544   0.01047   0.00359  -0.0927   0.5322   0.5526
   0.250   0.4828   0.01046   0.00358  -0.0928   0.5299   0.5563
   0.500   0.5114   0.01053   0.00362  -0.0931   0.5277   0.5602
   0.750   0.5400   0.01063   0.00370  -0.0933   0.5255   0.5646
   1.000   0.5683   0.01066   0.00374  -0.0935   0.5234   0.5689
   1.250   0.5964   0.01066   0.00379  -0.0937   0.5211   0.5730
   1.500   0.6247   0.01070   0.00385  -0.0938   0.5189   0.5772
   1.750   0.6531   0.01076   0.00389  -0.0941   0.5167   0.5817
   2.000   0.6815   0.01081   0.00394  -0.0943   0.5146   0.5860
   2.250   0.7099   0.01091   0.00405  -0.0945   0.5123   0.5907
   2.500   0.7382   0.01103   0.00418  -0.0948   0.5103   0.5960
   2.750   0.7660   0.01107   0.00428  -0.0949   0.5083   0.6012
   3.000   0.7937   0.01111   0.00440  -0.0950   0.5061   0.6070
   3.250   0.8217   0.01119   0.00450  -0.0951   0.5038   0.6138
   3.500   0.8495   0.01124   0.00460  -0.0953   0.5016   0.6202
   3.750   0.8777   0.01132   0.00471  -0.0955   0.4995   0.6276
   4.000   0.9058   0.01144   0.00484  -0.0957   0.4973   0.6349
   4.250   0.9336   0.01159   0.00503  -0.0959   0.4950   0.6442
   4.500   0.9602   0.01160   0.00518  -0.0958   0.4926   0.6540
   4.750   0.9872   0.01165   0.00532  -0.0958   0.4900   0.6651
   5.000   1.0144   0.01170   0.00545  -0.0959   0.4874   0.6779
   5.250   1.0417   0.01173   0.00557  -0.0959   0.4848   0.6933
   5.500   1.0689   0.01180   0.00571  -0.0960   0.4822   0.7126
   5.750   1.0951   0.01189   0.00593  -0.0958   0.4796   0.7393
   6.000   1.1196   0.01185   0.00612  -0.0953   0.4768   0.7785
   6.250   1.1415   0.01173   0.00630  -0.0941   0.4741   0.8521
   6.500   1.1762   0.01165   0.00641  -0.0956   0.4713   1.0000
   6.750   1.2036   0.01177   0.00655  -0.0958   0.4687   1.0000
   7.000   1.2314   0.01201   0.00676  -0.0960   0.4658   1.0000
   7.250   1.2562   0.01211   0.00697  -0.0957   0.4628   1.0000
   7.500   1.2819   0.01224   0.00717  -0.0955   0.4596   1.0000
   7.750   1.3080   0.01235   0.00733  -0.0954   0.4562   1.0000
   8.000   1.3344   0.01246   0.00740  -0.0953   0.4513   1.0000
   8.250   1.3545   0.01232   0.00739  -0.0940   0.4424   1.0000
   8.500   1.3751   0.01226   0.00735  -0.0928   0.4310   1.0000
   8.750   1.3956   0.01229   0.00740  -0.0916   0.4183   1.0000
   9.000   1.4161   0.01243   0.00755  -0.0905   0.4063   1.0000
   9.250   1.4357   0.01265   0.00779  -0.0893   0.3939   1.0000
   9.500   1.4536   0.01296   0.00810  -0.0878   0.3802   1.0000
   9.750   1.4681   0.01339   0.00850  -0.0858   0.3637   1.0000
  10.000   1.4772   0.01391   0.00899  -0.0828   0.3417   1.0000
  10.250   1.4774   0.01471   0.00968  -0.0785   0.3195   1.0000
  10.500   1.4731   0.01577   0.01063  -0.0738   0.2927   1.0000
  10.750   1.4623   0.01726   0.01197  -0.0688   0.2603   1.0000
  11.000   1.4435   0.01944   0.01399  -0.0638   0.2268   1.0000
  11.250   1.4196   0.02247   0.01687  -0.0595   0.1956   1.0000
  11.500   1.3998   0.02581   0.02011  -0.0566   0.1691   1.0000
  11.750   1.3786   0.02965   0.02385  -0.0543   0.1443   1.0000
  12.000   1.3583   0.03364   0.02776  -0.0525   0.1213   1.0000
  12.250   1.3392   0.03768   0.03171  -0.0510   0.1009   1.0000
  12.500   1.3203   0.04183   0.03578  -0.0497   0.0799   1.0000
  12.750   1.3030   0.04600   0.03986  -0.0487   0.0616   1.0000
  13.000   1.2887   0.05005   0.04385  -0.0481   0.0459   1.0000
  13.250   1.2777   0.05396   0.04771  -0.0477   0.0348   1.0000
  13.500   1.2695   0.05770   0.05144  -0.0474   0.0272   1.0000
  13.750   1.2665   0.06095   0.05473  -0.0473   0.0237   1.0000
  14.000   1.2623   0.06445   0.05826  -0.0473   0.0207   1.0000
  14.250   1.2631   0.06739   0.06127  -0.0474   0.0191   1.0000
  14.500   1.2593   0.07096   0.06489  -0.0477   0.0174   1.0000
  14.750   1.2566   0.07448   0.06847  -0.0479   0.0161   1.0000
  15.000   1.2585   0.07747   0.07154  -0.0483   0.0151   1.0000
  15.250   1.2583   0.08077   0.07491  -0.0487   0.0143   1.0000
  15.500   1.2554   0.08448   0.07866  -0.0493   0.0133   1.0000
  15.750   1.2492   0.08872   0.08298  -0.0500   0.0128   1.0000
<< Back to FX 66-S-171 AIRFOIL (fx66s171-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-S-171 AIRFOIL (fx66s171-il)