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

RAF 34 AIRFOIL (raf34-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 34 AIRFOIL (raf34-il)
Reynolds number: 100,000
Max Cl/Cd: 48.22 at α=8.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf34-il-100000.txt
Download as CSV file: xf-raf34-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 34 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
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.4741   0.13143   0.12610  -0.0062   1.0000   0.1176
 -11.750  -0.4825   0.12837   0.12309  -0.0095   1.0000   0.1221
 -11.500  -0.5242   0.12594   0.12081  -0.0178   1.0000   0.1239
 -11.250  -0.4717   0.12031   0.11507  -0.0110   1.0000   0.1278
 -11.000  -0.4632   0.11713   0.11189  -0.0113   1.0000   0.1325
 -10.750  -0.4832   0.11389   0.10875  -0.0161   1.0000   0.1380
 -10.500  -0.4969   0.10905   0.10400  -0.0196   1.0000   0.1401
 -10.250  -0.4668   0.10564   0.10053  -0.0164   1.0000   0.1430
 -10.000  -0.4587   0.10248   0.09738  -0.0165   1.0000   0.1476
  -9.750  -0.5042   0.09830   0.09337  -0.0248   1.0000   0.1541
  -9.500  -0.4854   0.09374   0.08882  -0.0229   1.0000   0.1566
  -9.250  -0.4618   0.09140   0.08645  -0.0204   1.0000   0.1620
  -9.000  -0.4961   0.08669   0.08189  -0.0265   1.0000   0.1689
  -8.750  -0.4903   0.08258   0.07783  -0.0261   1.0000   0.1731
  -8.500  -0.4761   0.08012   0.07537  -0.0247   1.0000   0.1793
  -8.250  -0.5403   0.07552   0.07087  -0.0286   1.0000   0.1847
  -8.000  -0.5182   0.07194   0.06736  -0.0270   1.0000   0.1900
  -7.750  -0.5200   0.06905   0.06450  -0.0262   1.0000   0.1963
  -7.500  -0.6176   0.05113   0.04559  -0.0272   1.0000   0.1126
  -7.250  -0.6522   0.04534   0.03886  -0.0193   1.0000   0.0941
  -7.000  -0.6599   0.04329   0.03676  -0.0147   1.0000   0.0933
  -6.750  -0.6664   0.04134   0.03469  -0.0104   1.0000   0.0922
  -6.500  -0.6631   0.03890   0.03200  -0.0077   0.9984   0.0909
  -6.250  -0.6293   0.03514   0.02768  -0.0101   0.9887   0.0896
  -6.000  -0.5926   0.03250   0.02451  -0.0126   0.9793   0.0914
  -5.750  -0.5535   0.03015   0.02166  -0.0150   0.9707   0.0929
  -5.500  -0.5155   0.02821   0.01928  -0.0171   0.9610   0.0941
  -5.250  -0.4688   0.02600   0.01688  -0.0209   0.9551   0.0969
  -5.000  -0.4298   0.02478   0.01564  -0.0233   0.9454   0.1025
  -4.750  -0.3785   0.02340   0.01401  -0.0275   0.9401   0.1088
  -4.500  -0.3403   0.02196   0.01275  -0.0296   0.9305   0.1166
  -4.250  -0.2948   0.02063   0.01150  -0.0330   0.9241   0.1335
  -4.000  -0.2663   0.01942   0.01054  -0.0331   0.9128   0.1652
  -3.750  -0.2409   0.01811   0.00977  -0.0328   0.9026   0.2409
  -3.500  -0.2165   0.01693   0.00926  -0.0323   0.8929   0.3581
  -3.250  -0.2044   0.01585   0.00927  -0.0289   0.8811   0.5650
  -3.000  -0.1441   0.01580   0.01001  -0.0315   0.8765   0.8079
  -2.750  -0.0861   0.01672   0.01077  -0.0344   0.8679   0.8680
  -2.500  -0.0277   0.01734   0.01112  -0.0381   0.8604   0.8955
  -2.250   0.0350   0.01790   0.01144  -0.0434   0.8504   0.9148
  -2.000   0.0949   0.01822   0.01154  -0.0485   0.8410   0.9347
  -1.750   0.1694   0.01836   0.01146  -0.0565   0.8310   0.9622
  -1.500   0.2320   0.01810   0.01104  -0.0633   0.8196   0.9836
  -1.250   0.2844   0.01758   0.01036  -0.0686   0.8085   0.9968
  -1.000   0.3097   0.01742   0.01009  -0.0687   0.7973   1.0000
  -0.750   0.3263   0.01744   0.01004  -0.0674   0.7851   1.0000
  -0.500   0.3435   0.01748   0.01000  -0.0659   0.7743   1.0000
  -0.250   0.3613   0.01748   0.00989  -0.0643   0.7646   1.0000
   0.000   0.3790   0.01757   0.00995  -0.0630   0.7529   1.0000
   0.250   0.3972   0.01764   0.00996  -0.0615   0.7429   1.0000
   0.500   0.4160   0.01768   0.00991  -0.0600   0.7332   1.0000
   0.750   0.4344   0.01781   0.01004  -0.0587   0.7220   1.0000
   1.000   0.4539   0.01788   0.01003  -0.0571   0.7130   1.0000
   1.250   0.4731   0.01799   0.01011  -0.0557   0.7025   1.0000
   1.500   0.4923   0.01815   0.01025  -0.0544   0.6922   1.0000
   1.750   0.5131   0.01816   0.01020  -0.0529   0.6837   1.0000
   2.000   0.5321   0.01839   0.01045  -0.0516   0.6724   1.0000
   2.250   0.5526   0.01851   0.01053  -0.0502   0.6633   1.0000
   2.500   0.5730   0.01863   0.01065  -0.0488   0.6535   1.0000
   2.750   0.5926   0.01888   0.01092  -0.0474   0.6429   1.0000
   3.000   0.6149   0.01890   0.01087  -0.0460   0.6350   1.0000
   3.250   0.6337   0.01922   0.01125  -0.0446   0.6234   1.0000
   3.500   0.6543   0.01940   0.01144  -0.0432   0.6134   1.0000
   3.750   0.6765   0.01944   0.01143  -0.0418   0.6040   1.0000
   4.000   0.6957   0.01966   0.01171  -0.0402   0.5919   1.0000
   4.250   0.7167   0.01978   0.01184  -0.0387   0.5810   1.0000
   4.500   0.7400   0.01975   0.01176  -0.0374   0.5714   1.0000
   4.750   0.7592   0.02000   0.01208  -0.0358   0.5592   1.0000
   5.000   0.7799   0.02017   0.01229  -0.0344   0.5481   1.0000
   5.250   0.8043   0.02009   0.01215  -0.0331   0.5385   1.0000
   5.500   0.8241   0.02022   0.01235  -0.0315   0.5255   1.0000
   5.750   0.8441   0.02040   0.01259  -0.0300   0.5130   1.0000
   6.000   0.8655   0.02056   0.01280  -0.0286   0.5018   1.0000
   6.250   0.8899   0.02056   0.01275  -0.0275   0.4919   1.0000
   6.500   0.9087   0.02088   0.01320  -0.0259   0.4793   1.0000
   6.750   0.9288   0.02113   0.01354  -0.0243   0.4668   1.0000
   7.000   0.9494   0.02123   0.01368  -0.0227   0.4528   1.0000
   7.250   0.9694   0.02119   0.01365  -0.0209   0.4362   1.0000
   7.500   0.9887   0.02111   0.01355  -0.0190   0.4177   1.0000
   7.750   1.0076   0.02108   0.01347  -0.0171   0.3987   1.0000
   8.000   1.0222   0.02130   0.01380  -0.0147   0.3779   1.0000
   8.250   1.0368   0.02150   0.01401  -0.0123   0.3559   1.0000
   8.500   1.0498   0.02183   0.01435  -0.0097   0.3330   1.0000
   8.750   1.0605   0.02225   0.01476  -0.0069   0.3079   1.0000
   9.000   1.0686   0.02284   0.01537  -0.0039   0.2802   1.0000
   9.250   1.0732   0.02361   0.01607  -0.0004   0.2503   1.0000
   9.500   1.0725   0.02464   0.01694   0.0036   0.2182   1.0000
   9.750   1.0656   0.02603   0.01812   0.0082   0.1857   1.0000
  10.000   1.0511   0.02765   0.01949   0.0137   0.1599   1.0000
  10.250   1.0373   0.02962   0.02125   0.0183   0.1379   1.0000
  10.500   1.0290   0.03171   0.02317   0.0217   0.1221   1.0000
  10.750   1.0275   0.03371   0.02505   0.0242   0.1096   1.0000
  11.000   1.0312   0.03552   0.02678   0.0263   0.1002   1.0000
  11.250   1.0371   0.03717   0.02840   0.0279   0.0926   1.0000
  11.500   1.0472   0.03878   0.03006   0.0294   0.0865   1.0000
  11.750   1.0571   0.04033   0.03159   0.0308   0.0813   1.0000
  12.000   1.0704   0.04198   0.03325   0.0320   0.0767   1.0000
  12.250   1.0811   0.04368   0.03507   0.0333   0.0729   1.0000
  12.500   1.0994   0.04524   0.03655   0.0342   0.0692   1.0000
  12.750   1.1169   0.04747   0.03892   0.0352   0.0666   1.0000
  13.000   1.1213   0.04975   0.04143   0.0364   0.0642   1.0000
  13.250   1.1277   0.05212   0.04398   0.0374   0.0624   1.0000
  13.500   1.1340   0.05446   0.04644   0.0382   0.0606   1.0000
  13.750   1.1454   0.05721   0.04926   0.0389   0.0592   1.0000
  14.000   1.1541   0.06126   0.05345   0.0395   0.0581   1.0000
  14.250   1.1391   0.06482   0.05731   0.0401   0.0579   1.0000
  14.500   1.1253   0.06867   0.06141   0.0403   0.0578   1.0000
  14.750   1.1079   0.07309   0.06609   0.0399   0.0578   1.0000
  15.000   1.0885   0.07794   0.07117   0.0390   0.0578   1.0000
  15.250   1.0635   0.08346   0.07693   0.0372   0.0578   1.0000
  15.500   1.0408   0.08943   0.08308   0.0350   0.0579   1.0000
  15.750   1.0163   0.09605   0.08989   0.0322   0.0581   1.0000
<< Back to RAF 34 AIRFOIL (raf34-il)

Polar data table (+)

Polar graphs


<< Back to RAF 34 AIRFOIL (raf34-il)