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NASA RC(4)-10 AIRFOIL (rc410-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NASA RC(4)-10 AIRFOIL (rc410-il)
Reynolds number: 50,000
Max Cl/Cd: 28.87 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rc410-il-50000.txt
Download as CSV file: xf-rc410-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA RC(4)-10 AIRFOIL                           
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5360   0.11525   0.10872   0.0067   1.0000   0.2024
  -8.750  -0.4979   0.10887   0.10228   0.0110   1.0000   0.2139
  -8.500  -0.5288   0.10759   0.10117   0.0044   1.0000   0.2186
  -8.250  -0.4970   0.10222   0.09578   0.0081   1.0000   0.2325
  -8.000  -0.4867   0.09811   0.09171   0.0082   1.0000   0.2419
  -7.750  -0.5184   0.09636   0.09014   0.0018   1.0000   0.2515
  -7.500  -0.4983   0.09208   0.08589   0.0041   1.0000   0.2670
  -7.250  -0.4854   0.08823   0.08209   0.0052   1.0000   0.2822
  -7.000  -0.4909   0.08520   0.07916   0.0035   1.0000   0.3006
  -6.750  -0.4727   0.08138   0.07538   0.0064   1.0000   0.3223
  -6.500  -0.4637   0.07856   0.07265   0.0088   1.0000   0.3554
  -6.250  -0.4608   0.07667   0.07085   0.0119   1.0000   0.4003
  -6.000  -0.4118   0.07378   0.06789   0.0205   1.0000   0.4742
  -5.750  -0.3739   0.07189   0.06598   0.0286   1.0000   0.5785
  -5.500  -0.3024   0.06788   0.06183   0.0345   1.0000   0.7299
  -4.750  -0.3118   0.05850   0.05276   0.0285   1.0000   0.6683
  -4.500  -0.4513   0.04758   0.04034  -0.0107   1.0000   0.2121
  -4.250  -0.4341   0.04412   0.03603  -0.0106   1.0000   0.1846
  -4.000  -0.4191   0.04136   0.03299  -0.0094   1.0000   0.1829
  -3.750  -0.4017   0.03880   0.03005  -0.0084   1.0000   0.1819
  -3.500  -0.3824   0.03685   0.02748  -0.0074   1.0000   0.1858
  -3.250  -0.3634   0.03450   0.02506  -0.0067   1.0000   0.1935
  -3.000  -0.3411   0.03304   0.02299  -0.0059   1.0000   0.2026
  -2.750  -0.3193   0.03117   0.02106  -0.0054   1.0000   0.2150
  -2.500  -0.2956   0.02965   0.01937  -0.0051   1.0000   0.2275
  -2.250  -0.2708   0.02842   0.01790  -0.0048   1.0000   0.2403
  -2.000  -0.2460   0.02747   0.01675  -0.0046   1.0000   0.2539
  -1.750  -0.2208   0.02671   0.01585  -0.0044   1.0000   0.2681
  -1.500  -0.1947   0.02593   0.01507  -0.0043   1.0000   0.2822
  -1.250  -0.1679   0.02537   0.01450  -0.0044   1.0000   0.2995
  -1.000  -0.1423   0.02489   0.01405  -0.0043   1.0000   0.3173
  -0.750  -0.1181   0.02436   0.01372  -0.0042   1.0000   0.3422
  -0.500  -0.0941   0.02374   0.01349  -0.0042   1.0000   0.3852
  -0.250  -0.0492   0.02204   0.01387  -0.0061   1.0000   1.0000
   0.000  -0.0298   0.02247   0.01390  -0.0051   1.0000   1.0000
   0.250  -0.0112   0.02292   0.01403  -0.0043   1.0000   1.0000
   0.500   0.0073   0.02340   0.01425  -0.0035   1.0000   1.0000
   0.750   0.0257   0.02390   0.01455  -0.0029   1.0000   1.0000
   1.000   0.1063   0.02526   0.01564  -0.0138   0.9744   1.0000
   1.250   0.2048   0.02631   0.01654  -0.0267   0.9356   1.0000
   1.500   0.2691   0.02688   0.01707  -0.0331   0.9115   1.0000
   1.750   0.3288   0.02725   0.01746  -0.0383   0.8876   1.0000
   2.000   0.4019   0.02717   0.01747  -0.0446   0.8611   1.0000
   2.250   0.4432   0.02709   0.01746  -0.0451   0.8335   1.0000
   2.500   0.4827   0.02672   0.01717  -0.0443   0.8043   1.0000
   2.750   0.5143   0.02609   0.01661  -0.0414   0.7719   1.0000
   3.000   0.5398   0.02516   0.01572  -0.0368   0.7355   1.0000
   3.250   0.5595   0.02394   0.01452  -0.0310   0.6890   1.0000
   3.500   0.5776   0.02242   0.01288  -0.0247   0.6261   1.0000
   3.750   0.5960   0.02146   0.01153  -0.0195   0.5492   1.0000
   4.000   0.6160   0.02150   0.01103  -0.0163   0.4869   1.0000
   4.250   0.6374   0.02208   0.01114  -0.0143   0.4461   1.0000
   4.500   0.6603   0.02294   0.01167  -0.0130   0.4174   1.0000
   4.750   0.6841   0.02390   0.01233  -0.0119   0.3956   1.0000
   5.000   0.7083   0.02495   0.01325  -0.0110   0.3770   1.0000
   5.250   0.7333   0.02609   0.01428  -0.0104   0.3628   1.0000
   5.500   0.7585   0.02729   0.01536  -0.0097   0.3501   1.0000
   5.750   0.7822   0.02856   0.01673  -0.0092   0.3378   1.0000
   6.000   0.8060   0.02998   0.01822  -0.0086   0.3276   1.0000
   6.250   0.8305   0.03144   0.01968  -0.0081   0.3191   1.0000
   6.500   0.8521   0.03313   0.02160  -0.0076   0.3109   1.0000
   6.750   0.8757   0.03471   0.02315  -0.0071   0.3024   1.0000
   7.000   0.8946   0.03676   0.02555  -0.0065   0.2957   1.0000
   7.250   0.9143   0.03881   0.02783  -0.0059   0.2898   1.0000
   7.500   0.9361   0.04090   0.02992  -0.0054   0.2838   1.0000
   7.750   0.9467   0.04364   0.03317  -0.0047   0.2781   1.0000
   8.000   0.9613   0.04625   0.03604  -0.0041   0.2730   1.0000
   8.250   0.9833   0.04869   0.03846  -0.0036   0.2686   1.0000
   8.500   0.9819   0.05284   0.04312  -0.0030   0.2661   1.0000
   8.750   0.9728   0.05776   0.04847  -0.0027   0.2640   1.0000
   9.000   0.9532   0.06383   0.05486  -0.0030   0.2635   1.0000
   9.250   0.9239   0.07129   0.06250  -0.0045   0.2655   1.0000
   9.500   0.9008   0.07875   0.07004  -0.0066   0.2690   1.0000
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