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FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: FX 66-H-60 (fx66h60-il)
Reynolds number: 200,000
Max Cl/Cd: 51.35 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx66h60-il-200000.txt
Download as CSV file: xf-fx66h60-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-H-60                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6093   0.10255   0.09933   0.0266   1.0000   0.0238
  -8.000  -0.6090   0.09892   0.09574   0.0235   1.0000   0.0239
  -7.750  -0.6096   0.09536   0.09220   0.0203   1.0000   0.0240
  -7.500  -0.6050   0.09155   0.08836   0.0175   1.0000   0.0240
  -7.250  -0.5980   0.08777   0.08454   0.0150   1.0000   0.0242
  -7.000  -0.5897   0.08393   0.08064   0.0130   1.0000   0.0242
  -6.750  -0.5799   0.08008   0.07671   0.0114   1.0000   0.0243
  -6.500  -0.5689   0.07618   0.07265   0.0102   1.0000   0.0244
  -6.250  -0.5640   0.06968   0.06610   0.0089   1.0000   0.0248
  -6.000  -0.5593   0.06377   0.06021   0.0086   1.0000   0.0257
  -5.750  -0.5486   0.05990   0.05632   0.0086   1.0000   0.0267
  -5.500  -0.5347   0.05632   0.05266   0.0084   1.0000   0.0278
  -5.250  -0.5184   0.05280   0.04902   0.0080   1.0000   0.0293
  -5.000  -0.4999   0.04932   0.04539   0.0078   1.0000   0.0311
  -4.750  -0.4788   0.04603   0.04188   0.0079   1.0000   0.0336
  -4.500  -0.4468   0.04576   0.04104   0.0091   1.0000   0.0368
  -4.250  -0.4332   0.03936   0.03449   0.0091   1.0000   0.0382
  -4.000  -0.4156   0.03588   0.03103   0.0093   1.0000   0.0406
  -3.750  -0.3934   0.03339   0.02836   0.0100   1.0000   0.0449
  -3.500  -0.3685   0.03081   0.02530   0.0112   1.0000   0.0521
  -3.250  -0.3455   0.02908   0.02324   0.0120   1.0000   0.0648
  -3.000  -0.3231   0.02687   0.02089   0.0125   1.0000   0.0785
  -2.750  -0.2998   0.02473   0.01863   0.0131   1.0000   0.0924
  -2.500  -0.2764   0.02257   0.01645   0.0135   1.0000   0.1081
  -2.000  -0.2087   0.01729   0.01008   0.0179   1.0000   0.0357
  -1.750  -0.1815   0.01526   0.00791   0.0187   1.0000   0.0333
  -1.500  -0.1546   0.01382   0.00640   0.0196   1.0000   0.0330
  -1.250  -0.1055   0.01247   0.00489   0.0156   0.8848   0.0382
  -1.000  -0.0732   0.01182   0.00390   0.0153   0.7654   0.0457
  -0.750   0.0404   0.00888   0.00329   0.0005   0.6622   1.0000
  -0.500   0.0623   0.00904   0.00307   0.0018   0.6134   1.0000
  -0.250   0.0848   0.00919   0.00286   0.0029   0.5757   1.0000
   0.000   0.1078   0.00931   0.00272   0.0039   0.5440   1.0000
   0.250   0.1311   0.00944   0.00262   0.0048   0.5175   1.0000
   0.500   0.1546   0.00957   0.00254   0.0056   0.4955   1.0000
   0.750   0.1784   0.00968   0.00250   0.0064   0.4750   1.0000
   1.000   0.2022   0.00982   0.00247   0.0071   0.4576   1.0000
   1.250   0.2260   0.00995   0.00247   0.0079   0.4417   1.0000
   1.500   0.2499   0.01010   0.00251   0.0086   0.4273   1.0000
   1.750   0.2738   0.01025   0.00256   0.0094   0.4141   1.0000
   2.000   0.2976   0.01042   0.00264   0.0102   0.4021   1.0000
   2.250   0.3215   0.01057   0.00274   0.0109   0.3903   1.0000
   2.500   0.3454   0.01075   0.00291   0.0116   0.3796   1.0000
   2.750   0.3692   0.01095   0.00308   0.0124   0.3701   1.0000
   3.000   0.3929   0.01119   0.00324   0.0132   0.3615   1.0000
   3.250   0.4168   0.01136   0.00346   0.0139   0.3521   1.0000
   3.500   0.4407   0.01161   0.00371   0.0147   0.3441   1.0000
   3.750   0.4646   0.01186   0.00401   0.0154   0.3369   1.0000
   4.000   0.4885   0.01212   0.00434   0.0162   0.3294   1.0000
   4.250   0.5124   0.01240   0.00462   0.0169   0.3226   1.0000
   4.500   0.5365   0.01269   0.00503   0.0176   0.3156   1.0000
   4.750   0.5606   0.01301   0.00539   0.0183   0.3098   1.0000
   5.000   0.5848   0.01335   0.00588   0.0190   0.3032   1.0000
   5.250   0.6086   0.01350   0.00614   0.0198   0.2926   1.0000
   5.500   0.6316   0.01311   0.00581   0.0207   0.2689   1.0000
   5.750   0.6557   0.01277   0.00562   0.0214   0.2372   1.0000
   6.000   0.6719   0.01509   0.00697   0.0225   0.0359   1.0000
   6.250   0.6926   0.01657   0.00870   0.0235   0.0257   1.0000
   6.500   0.7137   0.01786   0.01018   0.0246   0.0232   1.0000
   6.750   0.7339   0.01939   0.01183   0.0258   0.0218   1.0000
   7.000   0.7541   0.02106   0.01367   0.0270   0.0212   1.0000
   7.250   0.7748   0.02303   0.01577   0.0283   0.0213   1.0000
   7.500   0.7959   0.02538   0.01831   0.0295   0.0218   1.0000
   7.750   0.8162   0.02831   0.02154   0.0307   0.0229   1.0000
   8.000   0.8333   0.03243   0.02601   0.0317   0.0244   1.0000
   8.250   0.8502   0.03689   0.03073   0.0325   0.0257   1.0000
   8.500   0.8627   0.04099   0.03524   0.0330   0.0256   1.0000
  17.250   0.6614   0.20132   0.19816  -0.0463   0.0252   1.0000
  17.500   0.6649   0.20465   0.20150  -0.0483   0.0252   1.0000
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