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

NACA CYH (nacacyh-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA CYH (nacacyh-il)
Reynolds number: 50,000
Max Cl/Cd: 25.75 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nacacyh-il-50000.txt
Download as CSV file: xf-nacacyh-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA CYH                                        
  
 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  -0.4213   0.11394   0.10700  -0.0049   1.0000   0.2587
 -10.000  -0.4080   0.10974   0.10282  -0.0046   1.0000   0.2678
  -9.750  -0.4370   0.10960   0.10285  -0.0074   1.0000   0.2770
  -9.500  -0.4055   0.10408   0.09728  -0.0056   1.0000   0.2906
  -9.250  -0.3954   0.10024   0.09347  -0.0055   1.0000   0.2998
  -8.750  -0.4123   0.09599   0.08944  -0.0066   1.0000   0.3257
  -8.500  -0.3886   0.09154   0.08497  -0.0054   1.0000   0.3395
  -8.250  -0.3831   0.08850   0.08199  -0.0048   1.0000   0.3556
  -8.000  -0.3890   0.08625   0.07985  -0.0039   1.0000   0.3739
  -7.750  -0.3912   0.08360   0.07733  -0.0025   1.0000   0.3926
  -7.500  -0.3623   0.07960   0.07330  -0.0010   1.0000   0.4145
  -7.250  -0.3639   0.07742   0.07124   0.0011   1.0000   0.4378
  -6.750  -0.3717   0.07351   0.06757   0.0079   1.0000   0.4893
  -5.750  -0.4764   0.04956   0.04213  -0.0175   1.0000   0.1775
  -5.500  -0.4641   0.04531   0.03707  -0.0174   1.0000   0.1645
  -5.250  -0.4489   0.04266   0.03423  -0.0165   1.0000   0.1619
  -5.000  -0.4319   0.04015   0.03135  -0.0158   1.0000   0.1598
  -4.750  -0.4131   0.03804   0.02882  -0.0151   1.0000   0.1607
  -4.500  -0.3921   0.03620   0.02652  -0.0147   1.0000   0.1626
  -4.250  -0.3696   0.03453   0.02440  -0.0142   1.0000   0.1642
  -4.000  -0.3467   0.03298   0.02254  -0.0138   1.0000   0.1665
  -3.750  -0.3247   0.03181   0.02129  -0.0134   1.0000   0.1711
  -3.500  -0.3020   0.03095   0.02014  -0.0129   1.0000   0.1797
  -3.250  -0.2800   0.03009   0.01932  -0.0125   1.0000   0.1905
  -3.000  -0.2562   0.02930   0.01846  -0.0122   1.0000   0.2042
  -2.750  -0.2317   0.02868   0.01781  -0.0119   1.0000   0.2296
  -2.500  -0.2064   0.02794   0.01733  -0.0118   1.0000   0.2789
  -2.250  -0.1229   0.02454   0.01675  -0.0191   0.9950   1.0000
  -2.000  -0.0679   0.02543   0.01700  -0.0252   0.9820   1.0000
  -1.750  -0.0159   0.02621   0.01734  -0.0308   0.9664   1.0000
  -1.500   0.0345   0.02693   0.01773  -0.0358   0.9501   1.0000
  -1.250   0.0821   0.02759   0.01811  -0.0403   0.9340   1.0000
  -1.000   0.1271   0.02821   0.01852  -0.0441   0.9184   1.0000
  -0.750   0.1704   0.02880   0.01892  -0.0476   0.9031   1.0000
  -0.500   0.2128   0.02935   0.01934  -0.0507   0.8878   1.0000
  -0.250   0.2554   0.02987   0.01974  -0.0537   0.8723   1.0000
   0.000   0.2984   0.03034   0.02011  -0.0567   0.8565   1.0000
   0.250   0.3406   0.03076   0.02048  -0.0593   0.8405   1.0000
   0.500   0.3767   0.03118   0.02084  -0.0607   0.8239   1.0000
   0.750   0.4097   0.03161   0.02123  -0.0614   0.8070   1.0000
   1.000   0.4402   0.03205   0.02165  -0.0616   0.7900   1.0000
   1.250   0.4687   0.03251   0.02208  -0.0614   0.7729   1.0000
   1.500   0.4957   0.03298   0.02253  -0.0609   0.7560   1.0000
   1.750   0.5210   0.03349   0.02304  -0.0601   0.7393   1.0000
   2.000   0.5449   0.03405   0.02358  -0.0591   0.7230   1.0000
   2.250   0.5674   0.03467   0.02420  -0.0579   0.7073   1.0000
   2.500   0.5891   0.03535   0.02490  -0.0568   0.6924   1.0000
   2.750   0.6131   0.03591   0.02546  -0.0557   0.6787   1.0000
   3.000   0.6430   0.03610   0.02566  -0.0551   0.6668   1.0000
   3.250   0.6556   0.03735   0.02693  -0.0534   0.6523   1.0000
   3.500   0.6668   0.03876   0.02837  -0.0517   0.6385   1.0000
   3.750   0.6784   0.04021   0.02985  -0.0502   0.6254   1.0000
   4.000   0.6978   0.04119   0.03088  -0.0491   0.6144   1.0000
   4.250   0.7277   0.04133   0.03108  -0.0484   0.6043   1.0000
   4.500   0.7312   0.04348   0.03328  -0.0467   0.5911   1.0000
   4.750   0.7379   0.04545   0.03529  -0.0453   0.5787   1.0000
   5.000   0.7606   0.04612   0.03603  -0.0442   0.5678   1.0000
   5.250   0.7959   0.04560   0.03562  -0.0434   0.5575   1.0000
   5.500   0.7941   0.04834   0.03842  -0.0417   0.5441   1.0000
   5.750   0.7958   0.05085   0.04099  -0.0403   0.5312   1.0000
   6.000   0.8112   0.05210   0.04233  -0.0390   0.5191   1.0000
   6.250   0.8679   0.04915   0.03954  -0.0378   0.5081   1.0000
   6.500   0.9075   0.04761   0.03812  -0.0364   0.4945   1.0000
   6.750   0.9362   0.04699   0.03760  -0.0348   0.4792   1.0000
   7.000   0.9835   0.04445   0.03515  -0.0335   0.4629   1.0000
   7.250   0.9991   0.04505   0.03586  -0.0315   0.4456   1.0000
   7.500   1.0154   0.04564   0.03656  -0.0296   0.4280   1.0000
   7.750   1.0369   0.04580   0.03685  -0.0278   0.4099   1.0000
   8.000   1.0732   0.04469   0.03576  -0.0267   0.3901   1.0000
   8.250   1.1164   0.04336   0.03432  -0.0261   0.3694   1.0000
   8.500   1.1238   0.04510   0.03622  -0.0239   0.3513   1.0000
   8.750   1.1353   0.04672   0.03795  -0.0220   0.3343   1.0000
   9.000   1.1478   0.04842   0.03974  -0.0204   0.3188   1.0000
   9.250   1.1623   0.05000   0.04139  -0.0189   0.3044   1.0000
   9.500   1.1826   0.05125   0.04267  -0.0178   0.2908   1.0000
   9.750   1.1400   0.05717   0.04895  -0.0139   0.2857   1.0000
  10.000   1.1021   0.06292   0.05483  -0.0112   0.2813   1.0000
  10.250   1.0516   0.07118   0.06311  -0.0107   0.2785   1.0000
  10.500   0.8435   0.10849   0.09981  -0.0275   0.2895   1.0000
<< Back to NACA CYH (nacacyh-il)

Polar data table (+)

Polar graphs


<< Back to NACA CYH (nacacyh-il)