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

RAF 28 AIRFOIL (raf28-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: RAF 28 AIRFOIL (raf28-il)
Reynolds number: 1,000,000
Max Cl/Cd: 103.93 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf28-il-1000000.txt
Download as CSV file: xf-raf28-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 28 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.750  -0.8996   0.06434   0.06240  -0.0421   1.0000   0.0085
 -13.500  -0.9180   0.05537   0.05333  -0.0487   1.0000   0.0084
 -13.250  -0.9430   0.04438   0.04212  -0.0598   1.0000   0.0084
 -13.000  -0.5482   0.13094   0.12919  -0.0093   1.0000   0.0115
 -12.750  -0.9876   0.03470   0.03203  -0.0633   1.0000   0.0083
 -12.500  -0.9990   0.03281   0.03001  -0.0596   1.0000   0.0083
 -12.250  -1.0257   0.02877   0.02558  -0.0548   1.0000   0.0085
 -12.000  -1.0320   0.02608   0.02263  -0.0513   1.0000   0.0088
 -11.750  -1.0232   0.02474   0.02116  -0.0491   1.0000   0.0090
 -11.500  -1.0099   0.02379   0.02012  -0.0473   1.0000   0.0092
 -11.250  -0.9937   0.02312   0.01938  -0.0457   1.0000   0.0095
 -11.000  -0.9764   0.02254   0.01873  -0.0443   1.0000   0.0098
 -10.750  -0.9600   0.02182   0.01791  -0.0427   1.0000   0.0101
 -10.500  -0.9431   0.02112   0.01711  -0.0411   1.0000   0.0105
 -10.250  -0.9263   0.02038   0.01626  -0.0395   1.0000   0.0108
 -10.000  -0.9080   0.01980   0.01558  -0.0380   1.0000   0.0111
  -9.750  -0.8951   0.01845   0.01404  -0.0358   1.0000   0.0115
  -9.500  -0.8797   0.01746   0.01296  -0.0339   1.0000   0.0121
  -9.250  -0.8591   0.01718   0.01265  -0.0327   1.0000   0.0125
  -9.000  -0.8383   0.01696   0.01241  -0.0315   1.0000   0.0130
  -8.750  -0.8119   0.01653   0.01192  -0.0315   0.9994   0.0136
  -8.500  -0.7775   0.01597   0.01127  -0.0333   0.9976   0.0143
  -8.250  -0.7421   0.01573   0.01095  -0.0351   0.9958   0.0148
  -8.000  -0.7109   0.01451   0.00961  -0.0365   0.9939   0.0159
  -7.750  -0.6778   0.01443   0.00955  -0.0378   0.9913   0.0166
  -7.500  -0.6438   0.01424   0.00934  -0.0393   0.9886   0.0176
  -7.250  -0.6094   0.01390   0.00894  -0.0409   0.9862   0.0185
  -7.000  -0.5734   0.01374   0.00873  -0.0427   0.9842   0.0193
  -6.750  -0.5464   0.01256   0.00745  -0.0429   0.9790   0.0206
  -6.500  -0.5141   0.01215   0.00705  -0.0441   0.9745   0.0217
  -6.250  -0.4792   0.01184   0.00672  -0.0458   0.9713   0.0229
  -6.000  -0.4497   0.01146   0.00630  -0.0462   0.9621   0.0240
  -5.750  -0.4118   0.01112   0.00592  -0.0485   0.9572   0.0251
  -5.500  -0.3751   0.01105   0.00582  -0.0504   0.9463   0.0257
  -5.250  -0.3415   0.00979   0.00443  -0.0521   0.9314   0.0277
  -5.000  -0.3077   0.00935   0.00391  -0.0534   0.9108   0.0289
  -4.750  -0.2792   0.00911   0.00356  -0.0536   0.8888   0.0302
  -4.500  -0.2533   0.00890   0.00325  -0.0531   0.8719   0.0314
  -4.250  -0.2277   0.00870   0.00295  -0.0526   0.8589   0.0323
  -4.000  -0.2020   0.00853   0.00271  -0.0521   0.8486   0.0330
  -3.750  -0.1760   0.00841   0.00252  -0.0516   0.8395   0.0337
  -3.500  -0.1502   0.00822   0.00226  -0.0511   0.8312   0.0345
  -3.250  -0.1244   0.00802   0.00198  -0.0506   0.8237   0.0363
  -3.000  -0.0982   0.00786   0.00179  -0.0502   0.8160   0.0388
  -2.750  -0.0720   0.00773   0.00164  -0.0498   0.8090   0.0423
  -2.500  -0.0463   0.00751   0.00149  -0.0493   0.8013   0.0652
  -2.250  -0.0217   0.00718   0.00135  -0.0487   0.7941   0.1235
  -2.000   0.0017   0.00671   0.00122  -0.0479   0.7861   0.2201
  -1.750   0.0268   0.00647   0.00115  -0.0474   0.7785   0.2810
  -1.500   0.0524   0.00629   0.00109  -0.0469   0.7699   0.3265
  -1.250   0.0779   0.00610   0.00104  -0.0465   0.7608   0.3737
  -1.000   0.1027   0.00590   0.00100  -0.0458   0.7511   0.4336
  -0.750   0.1271   0.00565   0.00095  -0.0451   0.7398   0.5074
  -0.500   0.1503   0.00536   0.00092  -0.0442   0.7272   0.5935
  -0.250   0.1726   0.00511   0.00092  -0.0429   0.7131   0.6856
   0.250   0.2155   0.00474   0.00093  -0.0398   0.6787   0.8350
   0.500   0.2455   0.00464   0.00102  -0.0399   0.6606   0.9168
   0.750   0.2915   0.00478   0.00114  -0.0438   0.6440   0.9562
   1.000   0.3288   0.00493   0.00121  -0.0458   0.6250   0.9682
   1.250   0.3649   0.00510   0.00127  -0.0476   0.6059   0.9734
   1.500   0.4001   0.00522   0.00134  -0.0491   0.5918   0.9787
   1.750   0.4328   0.00533   0.00141  -0.0502   0.5795   0.9838
   2.000   0.4705   0.00545   0.00146  -0.0524   0.5624   0.9866
   2.250   0.5063   0.00557   0.00152  -0.0542   0.5408   0.9900
   2.500   0.5405   0.00570   0.00158  -0.0556   0.5222   0.9934
   2.750   0.5764   0.00585   0.00164  -0.0575   0.4961   0.9958
   3.000   0.6146   0.00603   0.00170  -0.0599   0.4642   0.9988
   3.250   0.6454   0.00621   0.00178  -0.0607   0.4355   1.0000
   3.500   0.6670   0.00642   0.00187  -0.0595   0.4043   1.0000
   3.750   0.6883   0.00668   0.00198  -0.0582   0.3690   1.0000
   4.000   0.7077   0.00708   0.00215  -0.0566   0.3138   1.0000
   4.250   0.7244   0.00769   0.00242  -0.0545   0.2376   1.0000
   4.500   0.7404   0.00838   0.00275  -0.0524   0.1627   1.0000
   4.750   0.7588   0.00889   0.00305  -0.0507   0.1190   1.0000
   5.000   0.7785   0.00930   0.00333  -0.0491   0.0890   1.0000
   5.250   0.7946   0.01001   0.00376  -0.0470   0.0383   1.0000
   5.500   0.8152   0.01037   0.00410  -0.0455   0.0312   1.0000
   5.750   0.8371   0.01063   0.00439  -0.0443   0.0286   1.0000
   6.000   0.8573   0.01103   0.00479  -0.0428   0.0250   1.0000
   6.250   0.8782   0.01138   0.00518  -0.0415   0.0229   1.0000
   6.500   0.8999   0.01166   0.00548  -0.0403   0.0212   1.0000
   6.750   0.9209   0.01199   0.00580  -0.0390   0.0194   1.0000
   7.000   0.9387   0.01259   0.00647  -0.0371   0.0176   1.0000
   7.250   0.9603   0.01288   0.00678  -0.0360   0.0168   1.0000
   7.500   0.9812   0.01323   0.00715  -0.0347   0.0158   1.0000
   7.750   1.0017   0.01360   0.00754  -0.0334   0.0149   1.0000
   8.000   1.0201   0.01413   0.00809  -0.0317   0.0140   1.0000
   8.250   1.0352   0.01490   0.00895  -0.0295   0.0131   1.0000
   8.500   1.0553   0.01528   0.00937  -0.0281   0.0126   1.0000
   8.750   1.0742   0.01575   0.00987  -0.0266   0.0121   1.0000
   9.000   1.0922   0.01627   0.01044  -0.0250   0.0116   1.0000
   9.250   1.1107   0.01673   0.01093  -0.0235   0.0111   1.0000
   9.500   1.1273   0.01731   0.01154  -0.0217   0.0106   1.0000
   9.750   1.1398   0.01818   0.01248  -0.0192   0.0102   1.0000
  10.000   1.1445   0.01964   0.01406  -0.0156   0.0097   1.0000
  10.250   1.1572   0.02034   0.01484  -0.0132   0.0096   1.0000
  10.500   1.1696   0.02093   0.01549  -0.0108   0.0093   1.0000
  10.750   1.1808   0.02164   0.01627  -0.0083   0.0091   1.0000
  11.000   1.1918   0.02240   0.01710  -0.0059   0.0088   1.0000
  11.250   1.2011   0.02334   0.01814  -0.0034   0.0086   1.0000
  11.500   1.2146   0.02390   0.01875  -0.0017   0.0082   1.0000
  11.750   1.2255   0.02472   0.01963   0.0002   0.0080   1.0000
  12.000   1.2319   0.02601   0.02103   0.0025   0.0080   1.0000
  12.250   1.2432   0.02680   0.02184   0.0040   0.0076   1.0000
  12.500   1.2452   0.02855   0.02369   0.0061   0.0074   1.0000
  12.750   1.2444   0.03070   0.02599   0.0082   0.0072   1.0000
  13.000   1.2340   0.03407   0.02959   0.0105   0.0070   1.0000
  13.250   1.2332   0.03647   0.03213   0.0119   0.0070   1.0000
  13.500   1.2371   0.03831   0.03410   0.0126   0.0069   1.0000
  13.750   1.2406   0.04030   0.03621   0.0132   0.0068   1.0000
  14.000   1.2369   0.04322   0.03928   0.0136   0.0068   1.0000
  14.250   1.2365   0.04588   0.04208   0.0136   0.0067   1.0000
  14.500   1.2279   0.04968   0.04605   0.0134   0.0067   1.0000
  14.750   1.2179   0.05387   0.05041   0.0126   0.0066   1.0000
  15.000   1.2081   0.05819   0.05489   0.0114   0.0066   1.0000
  15.250   1.2003   0.06250   0.05932   0.0098   0.0065   1.0000
  15.500   1.1862   0.06791   0.06490   0.0077   0.0064   1.0000
  15.750   1.1655   0.07476   0.07192   0.0046   0.0065   1.0000
  16.000   1.1411   0.08273   0.08008   0.0007   0.0065   1.0000
  16.250   1.1266   0.08937   0.08685  -0.0029   0.0064   1.0000
  16.500   1.0996   0.09866   0.09630  -0.0079   0.0066   1.0000
  16.750   1.0801   0.10705   0.10483  -0.0125   0.0065   1.0000
  17.000   1.0478   0.11874   0.11666  -0.0190   0.0066   1.0000
  17.250   1.0004   0.13502   0.13313  -0.0279   0.0068   1.0000
<< Back to RAF 28 AIRFOIL (raf28-il)

Polar data table (+)

Polar graphs


<< Back to RAF 28 AIRFOIL (raf28-il)