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

RAF 32 MOD AIRFOIL (raf32md-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 32 MOD AIRFOIL (raf32md-il)
Reynolds number: 100,000
Max Cl/Cd: 62.1 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf32md-il-100000.txt
Download as CSV file: xf-raf32md-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 32 MOD AIRFOIL                              
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3664   0.10733   0.10240  -0.0394   1.0000   0.1094
  -9.000  -0.3920   0.10679   0.10201  -0.0402   1.0000   0.1109
  -8.750  -0.4217   0.10648   0.10186  -0.0396   1.0000   0.1113
  -8.500  -0.3765   0.09977   0.09504  -0.0354   1.0000   0.1168
  -8.250  -0.3827   0.09798   0.09331  -0.0336   1.0000   0.1204
  -8.000  -0.3994   0.09664   0.09207  -0.0319   1.0000   0.1231
  -7.750  -0.4290   0.09616   0.09174  -0.0297   1.0000   0.1247
  -7.500  -0.3622   0.08680   0.08277  -0.0241   1.0000   0.1310
  -7.250  -0.3700   0.08481   0.08083  -0.0221   1.0000   0.1341
  -7.000  -0.4628   0.08951   0.08532  -0.0234   1.0000   0.1296
  -6.750  -0.4649   0.08742   0.08326  -0.0212   1.0000   0.1327
  -6.500  -0.4743   0.08499   0.08088  -0.0216   1.0000   0.1368
  -6.250  -0.4918   0.07987   0.07577  -0.0314   1.0000   0.1420
  -6.000  -0.4863   0.07789   0.07384  -0.0257   1.0000   0.1441
  -5.750  -0.4845   0.07383   0.06964  -0.0339   1.0000   0.1553
  -5.500  -0.4806   0.07024   0.06611  -0.0317   1.0000   0.1578
  -5.250  -0.4743   0.06837   0.06428  -0.0282   1.0000   0.1613
  -5.000  -0.4422   0.06280   0.05838  -0.0409   0.9953   0.1862
  -4.750  -0.4200   0.06037   0.05606  -0.0387   0.9907   0.1909
  -4.500  -0.3418   0.03448   0.02693  -0.0642   0.9901   0.0834
  -4.250  -0.3067   0.03179   0.02411  -0.0671   0.9856   0.0872
  -4.000  -0.2694   0.02968   0.02148  -0.0693   0.9807   0.0879
  -3.750  -0.2335   0.02812   0.01953  -0.0711   0.9751   0.0895
  -3.250  -0.1620   0.02612   0.01693  -0.0744   0.9637   0.0982
  -3.000  -0.1244   0.02522   0.01605  -0.0765   0.9587   0.1039
  -2.750  -0.0922   0.02465   0.01535  -0.0773   0.9520   0.1119
  -2.500  -0.0577   0.02394   0.01478  -0.0789   0.9460   0.1289
  -2.250  -0.0209   0.02279   0.01413  -0.0809   0.9414   0.2000
  -2.000   0.0065   0.02225   0.01420  -0.0814   0.9335   0.3456
  -1.750   0.0445   0.02206   0.01437  -0.0835   0.9286   0.4494
  -1.500   0.0661   0.02200   0.01463  -0.0825   0.9196   0.5359
  -1.250   0.1030   0.02192   0.01483  -0.0839   0.9145   0.6443
  -1.000   0.1230   0.02167   0.01495  -0.0819   0.9060   0.7530
  -0.750   0.1801   0.02123   0.01471  -0.0874   0.9017   1.0000
  -0.500   0.2116   0.02154   0.01477  -0.0887   0.8939   1.0000
  -0.250   0.2492   0.02176   0.01477  -0.0908   0.8872   1.0000
   0.000   0.2787   0.02208   0.01492  -0.0916   0.8791   1.0000
   0.250   0.3158   0.02227   0.01497  -0.0935   0.8726   1.0000
   0.500   0.3439   0.02259   0.01517  -0.0939   0.8643   1.0000
   0.750   0.3811   0.02274   0.01522  -0.0957   0.8579   1.0000
   1.000   0.4080   0.02308   0.01548  -0.0958   0.8494   1.0000
   1.250   0.4457   0.02317   0.01551  -0.0976   0.8432   1.0000
   1.500   0.4718   0.02351   0.01580  -0.0975   0.8345   1.0000
   1.750   0.5102   0.02353   0.01579  -0.0993   0.8285   1.0000
   2.000   0.5358   0.02388   0.01613  -0.0991   0.8197   1.0000
   2.250   0.5748   0.02381   0.01606  -0.1009   0.8139   1.0000
   2.500   0.6209   0.02354   0.01581  -0.1037   0.8100   1.0000
   2.750   0.6408   0.02397   0.01627  -0.1025   0.7993   1.0000
   3.000   0.6833   0.02374   0.01609  -0.1047   0.7946   1.0000
   3.250   0.7072   0.02402   0.01641  -0.1040   0.7846   1.0000
   3.500   0.7535   0.02353   0.01601  -0.1064   0.7808   1.0000
   3.750   0.7735   0.02392   0.01646  -0.1051   0.7695   1.0000
   4.000   0.7976   0.02419   0.01681  -0.1043   0.7594   1.0000
   4.250   0.8416   0.02360   0.01634  -0.1062   0.7539   1.0000
   4.500   0.8689   0.02350   0.01634  -0.1054   0.7424   1.0000
   4.750   0.9087   0.02258   0.01553  -0.1059   0.7311   1.0000
   5.000   0.9526   0.02146   0.01454  -0.1070   0.7200   1.0000
   5.250   0.9861   0.02069   0.01387  -0.1064   0.7045   1.0000
   5.500   1.0179   0.01974   0.01299  -0.1053   0.6838   1.0000
   5.750   1.0509   0.01887   0.01216  -0.1045   0.6615   1.0000
   6.000   1.0742   0.01861   0.01199  -0.1026   0.6387   1.0000
   6.250   1.0976   0.01841   0.01186  -0.1008   0.6140   1.0000
   6.500   1.1174   0.01829   0.01179  -0.0983   0.5833   1.0000
   6.750   1.1346   0.01827   0.01178  -0.0954   0.5447   1.0000
   7.000   1.1432   0.01851   0.01190  -0.0910   0.4866   1.0000
   7.250   1.1403   0.01939   0.01223  -0.0849   0.3810   1.0000
   7.500   1.1216   0.02156   0.01334  -0.0773   0.2432   1.0000
   7.750   1.1073   0.02395   0.01497  -0.0710   0.1706   1.0000
   8.000   1.1041   0.02592   0.01657  -0.0665   0.1384   1.0000
   8.250   1.1087   0.02756   0.01809  -0.0632   0.1165   1.0000
   8.500   1.1154   0.02922   0.01963  -0.0603   0.1015   1.0000
   8.750   1.1271   0.03100   0.02135  -0.0581   0.0893   1.0000
   9.000   1.1463   0.03292   0.02319  -0.0569   0.0788   1.0000
   9.250   1.1861   0.03562   0.02575  -0.0585   0.0702   1.0000
   9.500   1.2117   0.03749   0.02780  -0.0582   0.0639   1.0000
   9.750   1.2662   0.04202   0.03241  -0.0626   0.0588   1.0000
  10.000   1.2849   0.04452   0.03533  -0.0611   0.0570   1.0000
  10.250   1.2980   0.04697   0.03816  -0.0590   0.0546   1.0000
  10.500   1.3103   0.04965   0.04113  -0.0571   0.0526   1.0000
  10.750   1.3188   0.05293   0.04479  -0.0548   0.0521   1.0000
  11.000   1.3193   0.05641   0.04869  -0.0516   0.0521   1.0000
  11.250   1.3117   0.05993   0.05262  -0.0476   0.0526   1.0000
  11.500   1.2984   0.06357   0.05665  -0.0434   0.0533   1.0000
  11.750   1.2818   0.06734   0.06076  -0.0396   0.0541   1.0000
  12.000   1.2632   0.07136   0.06508  -0.0365   0.0549   1.0000
  12.250   1.2436   0.07563   0.06961  -0.0341   0.0557   1.0000
  12.500   1.2231   0.08022   0.07443  -0.0326   0.0565   1.0000
  12.750   1.2018   0.08518   0.07960  -0.0318   0.0571   1.0000
  13.000   1.1812   0.09062   0.08521  -0.0320   0.0578   1.0000
  13.250   1.1649   0.09670   0.09141  -0.0327   0.0587   1.0000
  13.500   1.1544   0.10184   0.09674  -0.0336   0.0607   1.0000
<< Back to RAF 32 MOD AIRFOIL (raf32md-il)

Polar data table (+)

Polar graphs


<< Back to RAF 32 MOD AIRFOIL (raf32md-il)