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NACA 65(1)-212 (naca651212-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-212 (naca651212-il)
Reynolds number: 200,000
Max Cl/Cd: 55.93 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651212-il-200000.txt
Download as CSV file: xf-naca651212-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-212                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5424   0.08422   0.08068  -0.0459   1.0000   0.0588
 -10.000  -0.5642   0.07758   0.07402  -0.0514   1.0000   0.0590
  -9.750  -0.5884   0.07283   0.06920  -0.0540   1.0000   0.0592
  -9.500  -0.6150   0.06975   0.06604  -0.0535   1.0000   0.0595
  -9.250  -0.6465   0.06822   0.06428  -0.0508   1.0000   0.0602
  -9.000  -0.6632   0.06705   0.06287  -0.0477   1.0000   0.0606
  -8.750  -0.6772   0.06552   0.06117  -0.0438   1.0000   0.0607
  -8.500  -0.6766   0.05684   0.05277  -0.0437   1.0000   0.0628
  -8.250  -0.6814   0.05507   0.05106  -0.0400   1.0000   0.0639
  -8.000  -0.6901   0.05364   0.04962  -0.0359   1.0000   0.0651
  -7.750  -0.6835   0.05141   0.04726  -0.0351   0.9976   0.0690
  -7.500  -0.6654   0.04691   0.04229  -0.0382   0.9905   0.0769
  -7.250  -0.6366   0.04398   0.03926  -0.0407   0.9858   0.0823
  -7.000  -0.6136   0.04049   0.03543  -0.0429   0.9797   0.0918
  -6.750  -0.5792   0.03199   0.02536  -0.0410   0.9740   0.0477
  -6.500  -0.5450   0.02848   0.02148  -0.0430   0.9711   0.0470
  -6.250  -0.5148   0.02584   0.01854  -0.0438   0.9654   0.0464
  -6.000  -0.4793   0.02354   0.01595  -0.0454   0.9615   0.0459
  -5.750  -0.4405   0.02174   0.01393  -0.0475   0.9590   0.0463
  -5.500  -0.4001   0.02066   0.01267  -0.0500   0.9567   0.0482
  -5.250  -0.3721   0.01891   0.01092  -0.0502   0.9501   0.0504
  -5.000  -0.3385   0.01769   0.00973  -0.0516   0.9456   0.0529
  -4.750  -0.3036   0.01675   0.00878  -0.0533   0.9420   0.0563
  -4.500  -0.2795   0.01626   0.00823  -0.0528   0.9333   0.0605
  -4.250  -0.2515   0.01525   0.00726  -0.0533   0.9282   0.0682
  -4.000  -0.2297   0.01467   0.00664  -0.0524   0.9194   0.0783
  -3.750  -0.2112   0.01290   0.00573  -0.0517   0.9131   0.2445
  -3.500  -0.2063   0.01144   0.00585  -0.0479   0.9032   0.6196
  -3.250  -0.1795   0.01158   0.00603  -0.0469   0.8985   0.6712
  -3.000  -0.1566   0.01181   0.00624  -0.0455   0.8908   0.6996
  -2.750  -0.1318   0.01212   0.00658  -0.0438   0.8854   0.7288
  -2.500  -0.1109   0.01270   0.00722  -0.0410   0.8793   0.7585
  -2.250  -0.0890   0.01312   0.00765  -0.0386   0.8730   0.7783
  -2.000  -0.0633   0.01326   0.00774  -0.0375   0.8687   0.7891
  -1.750  -0.0384   0.01325   0.00764  -0.0372   0.8615   0.7975
  -1.500  -0.0118   0.01322   0.00757  -0.0367   0.8563   0.8018
  -1.250   0.0156   0.01318   0.00746  -0.0365   0.8522   0.8074
  -1.000   0.0409   0.01316   0.00740  -0.0364   0.8451   0.8139
  -0.750   0.0676   0.01313   0.00734  -0.0360   0.8404   0.8183
  -0.500   0.0943   0.01312   0.00729  -0.0359   0.8356   0.8245
  -0.250   0.1199   0.01313   0.00728  -0.0357   0.8293   0.8301
   0.000   0.1468   0.01310   0.00724  -0.0353   0.8250   0.8351
   0.250   0.1735   0.01312   0.00724  -0.0355   0.8199   0.8422
   0.500   0.1981   0.01316   0.00731  -0.0348   0.8140   0.8470
   0.750   0.2253   0.01314   0.00729  -0.0346   0.8099   0.8530
   1.000   0.2511   0.01319   0.00736  -0.0344   0.8048   0.8594
   1.250   0.2755   0.01326   0.00748  -0.0338   0.7990   0.8652
   1.500   0.3032   0.01325   0.00746  -0.0338   0.7949   0.8726
   1.750   0.3273   0.01332   0.00760  -0.0331   0.7897   0.8780
   2.000   0.3521   0.01340   0.00774  -0.0327   0.7839   0.8859
   2.250   0.3783   0.01337   0.00775  -0.0321   0.7797   0.8919
   2.500   0.4021   0.01345   0.00791  -0.0315   0.7733   0.9000
   2.750   0.4272   0.01340   0.00791  -0.0307   0.7669   0.9068
   3.000   0.4529   0.01334   0.00792  -0.0300   0.7608   0.9155
   3.250   0.4786   0.01307   0.00770  -0.0290   0.7505   0.9226
   3.500   0.5049   0.01250   0.00709  -0.0275   0.7337   0.9311
   3.750   0.5300   0.01185   0.00643  -0.0257   0.7066   0.9387
   4.000   0.5559   0.01145   0.00604  -0.0247   0.6810   0.9477
   4.250   0.5852   0.01123   0.00583  -0.0245   0.6578   0.9566
   4.500   0.6175   0.01104   0.00563  -0.0251   0.6152   0.9638
   4.750   0.6397   0.01150   0.00541  -0.0239   0.4400   0.9740
   5.000   0.6441   0.01526   0.00705  -0.0227   0.0812   0.9882
   5.250   0.6780   0.01641   0.00817  -0.0249   0.0630   0.9977
   5.500   0.6847   0.01686   0.00864  -0.0218   0.0582   1.0000
   5.750   0.6881   0.01773   0.00946  -0.0183   0.0536   1.0000
   6.000   0.7063   0.01846   0.01024  -0.0171   0.0507   1.0000
   6.250   0.7261   0.01938   0.01117  -0.0161   0.0480   1.0000
   6.500   0.7476   0.02043   0.01221  -0.0154   0.0459   1.0000
   6.750   0.7716   0.02177   0.01350  -0.0152   0.0439   1.0000
   7.000   0.8002   0.02375   0.01551  -0.0156   0.0417   1.0000
   7.250   0.8277   0.02504   0.01693  -0.0156   0.0406   1.0000
   7.500   0.8563   0.02679   0.01884  -0.0157   0.0400   1.0000
   7.750   0.8841   0.02894   0.02121  -0.0157   0.0399   1.0000
   8.000   0.9097   0.03157   0.02412  -0.0154   0.0403   1.0000
   8.250   0.9316   0.03413   0.02698  -0.0146   0.0403   1.0000
   8.500   0.9503   0.03680   0.02997  -0.0135   0.0401   1.0000
   8.750   0.9737   0.04149   0.03479  -0.0134   0.0429   1.0000
  11.000   0.7424   0.07904   0.07589   0.0082   0.0743   1.0000
  11.250   0.7070   0.08775   0.08467   0.0039   0.0753   1.0000
  11.500   0.6706   0.09873   0.09569  -0.0031   0.0757   1.0000
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