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NACA 66(1)-212 (naca661212-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 66(1)-212 (naca661212-il)
Reynolds number: 100,000
Max Cl/Cd: 37.64 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca661212-il-100000.txt
Download as CSV file: xf-naca661212-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66(1)-212                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4800   0.09883   0.09386  -0.0451   1.0000   0.1295
 -10.000  -0.5165   0.09384   0.08903  -0.0515   1.0000   0.1332
  -9.750  -0.5039   0.09046   0.08568  -0.0487   1.0000   0.1374
  -9.500  -0.5033   0.08778   0.08305  -0.0476   1.0000   0.1435
  -9.250  -0.5382   0.08402   0.07943  -0.0488   1.0000   0.1458
  -9.000  -0.5718   0.08212   0.07762  -0.0456   1.0000   0.1459
  -8.750  -0.6070   0.08088   0.07647  -0.0406   1.0000   0.1456
  -8.500  -0.6409   0.07962   0.07526  -0.0353   1.0000   0.1455
  -8.250  -0.6759   0.07800   0.07361  -0.0309   1.0000   0.1465
  -8.000  -0.7180   0.07692   0.07236  -0.0263   1.0000   0.1480
  -6.750  -0.7349   0.06236   0.05773  -0.0131   1.0000   0.1866
  -6.500  -0.7369   0.05981   0.05513  -0.0104   1.0000   0.2009
  -5.250  -0.6719   0.03689   0.02883  -0.0022   1.0000   0.0856
  -5.000  -0.6489   0.03487   0.02593   0.0002   1.0000   0.0765
  -4.750  -0.6278   0.03236   0.02317   0.0013   1.0000   0.0753
  -4.500  -0.6066   0.03099   0.02144   0.0026   1.0000   0.0764
  -4.250  -0.5847   0.02959   0.01976   0.0038   1.0000   0.0769
  -4.000  -0.5620   0.02814   0.01811   0.0049   1.0000   0.0770
  -3.750  -0.5392   0.02656   0.01644   0.0058   1.0000   0.0779
  -3.500  -0.5175   0.02517   0.01513   0.0068   1.0000   0.0811
  -3.250  -0.4968   0.02438   0.01434   0.0078   1.0000   0.0866
  -3.000  -0.4759   0.02368   0.01359   0.0090   1.0000   0.0904
  -2.750  -0.4569   0.02271   0.01274   0.0103   1.0000   0.0953
  -2.500  -0.4372   0.02218   0.01216   0.0115   1.0000   0.1050
  -2.250  -0.4176   0.02151   0.01157   0.0126   1.0000   0.1208
  -2.000  -0.4033   0.01896   0.01089   0.0138   1.0000   0.4331
  -1.750  -0.4203   0.01935   0.01289   0.0272   1.0000   0.8245
  -1.500  -0.1148   0.02659   0.01930  -0.0137   1.0000   0.9982
  -1.250  -0.0976   0.02656   0.01917  -0.0134   1.0000   1.0000
  -1.000  -0.0769   0.02649   0.01901  -0.0136   0.9975   1.0000
  -0.750  -0.0505   0.02655   0.01898  -0.0150   0.9939   1.0000
  -0.500  -0.0290   0.02652   0.01888  -0.0154   0.9901   1.0000
  -0.250  -0.0050   0.02662   0.01892  -0.0162   0.9862   1.0000
   0.000   0.0207   0.02680   0.01906  -0.0173   0.9822   1.0000
   0.250   0.0392   0.02684   0.01907  -0.0170   0.9775   1.0000
   0.500   0.0655   0.02708   0.01927  -0.0182   0.9727   1.0000
   0.750   0.0867   0.02725   0.01944  -0.0184   0.9674   1.0000
   1.000   0.1081   0.02742   0.01961  -0.0186   0.9614   1.0000
   1.250   0.1374   0.02781   0.02001  -0.0202   0.9558   1.0000
   1.500   0.1520   0.02790   0.02012  -0.0191   0.9485   1.0000
   1.750   0.1865   0.02839   0.02064  -0.0216   0.9423   1.0000
   2.000   0.1981   0.02847   0.02075  -0.0198   0.9335   1.0000
   2.250   0.2257   0.02885   0.02119  -0.0209   0.9258   1.0000
   2.500   0.2489   0.02911   0.02150  -0.0212   0.9167   1.0000
   2.750   0.2672   0.02934   0.02180  -0.0204   0.9064   1.0000
   3.000   0.3025   0.02974   0.02231  -0.0226   0.8973   1.0000
   3.250   0.3292   0.02995   0.02262  -0.0232   0.8862   1.0000
   3.500   0.3491   0.03010   0.02286  -0.0224   0.8735   1.0000
   3.750   0.3731   0.03025   0.02315  -0.0222   0.8609   1.0000
   4.000   0.4032   0.03034   0.02339  -0.0228   0.8481   1.0000
   4.250   0.6914   0.01837   0.00967  -0.0452   0.2615   1.0000
   4.500   0.6814   0.02039   0.01062  -0.0391   0.1272   1.0000
   4.750   0.6846   0.02132   0.01141  -0.0349   0.1095   1.0000
   5.000   0.6884   0.02207   0.01215  -0.0309   0.0987   1.0000
   5.250   0.6925   0.02278   0.01288  -0.0268   0.0916   1.0000
   5.500   0.6975   0.02354   0.01361  -0.0230   0.0867   1.0000
   5.750   0.7110   0.02481   0.01475  -0.0206   0.0824   1.0000
   6.000   0.7284   0.02565   0.01570  -0.0187   0.0779   1.0000
   6.250   0.7498   0.02678   0.01681  -0.0177   0.0732   1.0000
   6.500   0.7992   0.02993   0.01984  -0.0214   0.0700   1.0000
   6.750   0.8222   0.03160   0.02173  -0.0203   0.0694   1.0000
   7.000   0.8384   0.03307   0.02350  -0.0180   0.0687   1.0000
   7.250   0.8490   0.03437   0.02511  -0.0149   0.0671   1.0000
   7.500   0.8613   0.03624   0.02727  -0.0121   0.0669   1.0000
   7.750   0.8736   0.03873   0.03003  -0.0094   0.0679   1.0000
   8.000   0.8884   0.04239   0.03388  -0.0076   0.0695   1.0000
   8.250   0.8916   0.04410   0.03645  -0.0021   0.0766   1.0000
   8.500   0.8994   0.05085   0.04421   0.0023   0.1065   1.0000
  10.750   0.7250   0.09928   0.09462   0.0092   0.1807   1.0000
  11.000   0.6943   0.10563   0.10087   0.0035   0.1710   1.0000
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