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

RAF 69 AIRFOIL (raf69-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: RAF 69 AIRFOIL (raf69-il)
Reynolds number: 100,000
Max Cl/Cd: 40.09 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf69-il-100000-n5.txt
Download as CSV file: xf-raf69-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 69 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -16.750  -0.8093   0.08086   0.07371  -0.0754   1.0000   0.1050
 -16.500  -0.7876   0.08120   0.07417  -0.0744   1.0000   0.1069
 -16.250  -0.7942   0.07743   0.07036  -0.0757   1.0000   0.1089
 -16.000  -0.8154   0.07177   0.06456  -0.0780   1.0000   0.1109
 -15.750  -0.8377   0.06626   0.05888  -0.0800   1.0000   0.1130
 -15.500  -0.8590   0.06124   0.05363  -0.0815   1.0000   0.1151
 -15.250  -0.8629   0.05845   0.05080  -0.0817   1.0000   0.1171
 -15.000  -0.8600   0.05658   0.04895  -0.0815   1.0000   0.1193
 -14.750  -0.8630   0.05415   0.04649  -0.0814   1.0000   0.1217
 -14.500  -0.8687   0.05152   0.04377  -0.0812   1.0000   0.1242
 -14.250  -0.8760   0.04887   0.04098  -0.0808   1.0000   0.1268
 -14.000  -0.8794   0.04671   0.03874  -0.0801   1.0000   0.1295
 -13.750  -0.8777   0.04520   0.03729  -0.0790   1.0000   0.1320
 -13.500  -0.8793   0.04354   0.03561  -0.0778   1.0000   0.1347
 -13.250  -0.8826   0.04185   0.03386  -0.0762   1.0000   0.1378
 -13.000  -0.8881   0.04019   0.03207  -0.0743   1.0000   0.1411
 -12.750  -0.8904   0.03896   0.03087  -0.0720   1.0000   0.1439
 -12.500  -0.8951   0.03792   0.02987  -0.0693   1.0000   0.1465
 -12.250  -0.9045   0.03696   0.02891  -0.0657   1.0000   0.1493
 -12.000  -0.9188   0.03614   0.02806  -0.0612   1.0000   0.1519
 -11.750  -0.9083   0.03488   0.02664  -0.0612   0.9946   0.1569
 -11.500  -0.8856   0.03380   0.02560  -0.0628   0.9857   0.1618
 -11.250  -0.8629   0.03271   0.02441  -0.0644   0.9776   0.1679
 -11.000  -0.8406   0.03169   0.02329  -0.0655   0.9689   0.1738
 -10.750  -0.8189   0.03085   0.02244  -0.0663   0.9603   0.1796
 -10.500  -0.7939   0.02995   0.02136  -0.0676   0.9531   0.1867
 -10.250  -0.7719   0.02923   0.02066  -0.0679   0.9445   0.1924
 -10.000  -0.7454   0.02848   0.01983  -0.0691   0.9379   0.1997
  -9.750  -0.7240   0.02783   0.01910  -0.0690   0.9292   0.2062
  -9.500  -0.6991   0.02722   0.01846  -0.0695   0.9218   0.2129
  -9.250  -0.6797   0.02666   0.01774  -0.0688   0.9128   0.2202
  -9.000  -0.6561   0.02613   0.01726  -0.0687   0.9051   0.2265
  -8.750  -0.6314   0.02560   0.01663  -0.0689   0.8984   0.2343
  -8.500  -0.6132   0.02519   0.01620  -0.0676   0.8887   0.2407
  -8.250  -0.5861   0.02470   0.01567  -0.0680   0.8828   0.2486
  -8.000  -0.5704   0.02438   0.01525  -0.0662   0.8733   0.2555
  -7.750  -0.5469   0.02398   0.01488  -0.0658   0.8663   0.2625
  -7.500  -0.5225   0.02358   0.01436  -0.0655   0.8601   0.2709
  -7.250  -0.5064   0.02332   0.01414  -0.0636   0.8508   0.2772
  -7.000  -0.4818   0.02295   0.01372  -0.0633   0.8446   0.2856
  -6.750  -0.4616   0.02266   0.01340  -0.0621   0.8374   0.2932
  -6.500  -0.4430   0.02241   0.01316  -0.0605   0.8293   0.3009
  -6.250  -0.4172   0.02206   0.01273  -0.0603   0.8236   0.3101
  -6.000  -0.3994   0.02185   0.01257  -0.0586   0.8158   0.3179
  -5.750  -0.3792   0.02162   0.01225  -0.0573   0.8083   0.3276
  -5.500  -0.3528   0.02128   0.01196  -0.0570   0.8029   0.3368
  -5.000  -0.3150   0.02090   0.01160  -0.0538   0.7875   0.3564
  -4.750  -0.2879   0.02060   0.01125  -0.0536   0.7824   0.3682
  -4.500  -0.2731   0.02052   0.01121  -0.0512   0.7734   0.3785
  -4.250  -0.2501   0.02030   0.01100  -0.0502   0.7668   0.3904
  -3.750  -0.2086   0.02001   0.01074  -0.0475   0.7528   0.4154
  -3.500  -0.1847   0.01982   0.01055  -0.0467   0.7466   0.4289
  -3.250  -0.1590   0.01964   0.01034  -0.0461   0.7413   0.4438
  -3.000  -0.1431   0.01962   0.01040  -0.0439   0.7328   0.4567
  -2.750  -0.1185   0.01947   0.01026  -0.0431   0.7267   0.4708
  -2.500  -0.0928   0.01932   0.01010  -0.0425   0.7213   0.4861
  -2.250  -0.0764   0.01934   0.01015  -0.0404   0.7129   0.5002
  -2.000  -0.0511   0.01921   0.01006  -0.0397   0.7068   0.5149
  -1.750  -0.0254   0.01911   0.00997  -0.0390   0.7012   0.5298
  -1.500  -0.0083   0.01914   0.01004  -0.0370   0.6929   0.5444
  -1.250   0.0176   0.01903   0.00994  -0.0364   0.6868   0.5598
  -1.000   0.0420   0.01899   0.00994  -0.0355   0.6805   0.5744
  -0.750   0.0610   0.01903   0.01004  -0.0337   0.6725   0.5888
  -0.500   0.0879   0.01894   0.00994  -0.0332   0.6664   0.6046
  -0.250   0.1091   0.01898   0.01001  -0.0318   0.6593   0.6200
   0.000   0.1308   0.01902   0.01012  -0.0304   0.6515   0.6345
   0.250   0.1596   0.01896   0.01006  -0.0302   0.6458   0.6497
   0.500   0.1770   0.01907   0.01022  -0.0281   0.6375   0.6647
   0.750   0.2006   0.01910   0.01026  -0.0271   0.6302   0.6797
   1.000   0.2298   0.01908   0.01023  -0.0269   0.6244   0.6937
   1.250   0.2455   0.01924   0.01046  -0.0246   0.6153   0.7076
   1.500   0.2710   0.01922   0.01040  -0.0238   0.6086   0.7226
   1.750   0.2926   0.01932   0.01055  -0.0224   0.6008   0.7356
   2.000   0.3145   0.01939   0.01064  -0.0211   0.5928   0.7487
   2.250   0.3431   0.01934   0.01053  -0.0209   0.5868   0.7629
   2.500   0.3582   0.01952   0.01079  -0.0185   0.5772   0.7760
   2.750   0.3863   0.01953   0.01079  -0.0182   0.5702   0.7884
   3.000   0.4053   0.01964   0.01092  -0.0165   0.5619   0.8020
   3.250   0.4306   0.01970   0.01101  -0.0158   0.5536   0.8145
   3.500   0.4582   0.01978   0.01108  -0.0156   0.5460   0.8264
   3.750   0.4779   0.01990   0.01124  -0.0140   0.5370   0.8401
   4.000   0.5127   0.01996   0.01127  -0.0151   0.5294   0.8507
   4.250   0.5342   0.02018   0.01156  -0.0140   0.5197   0.8635
   4.500   0.5694   0.02025   0.01157  -0.0152   0.5122   0.8747
   4.750   0.5951   0.02056   0.01197  -0.0151   0.5025   0.8863
   5.000   0.6268   0.02068   0.01203  -0.0158   0.4947   0.8985
   5.250   0.6599   0.02104   0.01246  -0.0171   0.4850   0.9081
   5.500   0.6938   0.02122   0.01260  -0.0183   0.4768   0.9189
   5.750   0.7268   0.02159   0.01301  -0.0198   0.4673   0.9291
   6.000   0.7635   0.02185   0.01323  -0.0217   0.4585   0.9383
   6.250   0.7945   0.02223   0.01365  -0.0228   0.4492   0.9495
   6.500   0.8323   0.02254   0.01392  -0.0251   0.4400   0.9574
   6.750   0.8631   0.02295   0.01437  -0.0263   0.4305   0.9687
   7.000   0.8967   0.02326   0.01464  -0.0280   0.4211   0.9795
   7.250   0.9282   0.02369   0.01513  -0.0296   0.4108   0.9909
   7.500   0.9629   0.02402   0.01538  -0.0316   0.4014   1.0000
   7.750   0.9428   0.02418   0.01556  -0.0239   0.3953   1.0000
   8.000   0.9360   0.02444   0.01579  -0.0184   0.3881   1.0000
   8.250   0.9412   0.02472   0.01598  -0.0148   0.3810   1.0000
   8.500   0.9382   0.02524   0.01654  -0.0103   0.3728   1.0000
   8.750   0.9463   0.02561   0.01682  -0.0074   0.3651   1.0000
   9.000   0.9474   0.02626   0.01751  -0.0038   0.3568   1.0000
   9.250   0.9542   0.02681   0.01802  -0.0011   0.3489   1.0000
   9.500   0.9596   0.02749   0.01871   0.0017   0.3409   1.0000
   9.750   0.9650   0.02821   0.01943   0.0043   0.3327   1.0000
  10.000   0.9733   0.02891   0.02010   0.0065   0.3251   1.0000
  10.250   0.9778   0.02982   0.02105   0.0090   0.3169   1.0000
  10.500   0.9881   0.03052   0.02168   0.0108   0.3097   1.0000
  10.750   0.9913   0.03165   0.02288   0.0130   0.3015   1.0000
  11.000   1.0007   0.03246   0.02363   0.0147   0.2943   1.0000
  11.250   1.0051   0.03366   0.02490   0.0165   0.2867   1.0000
  11.500   1.0121   0.03472   0.02594   0.0181   0.2795   1.0000
  11.750   1.0192   0.03587   0.02710   0.0196   0.2726   1.0000
  12.000   1.0237   0.03720   0.02848   0.0211   0.2654   1.0000
  12.250   1.0339   0.03818   0.02938   0.0223   0.2593   1.0000
  12.500   1.0360   0.03984   0.03114   0.0237   0.2523   1.0000
  12.750   1.0437   0.04108   0.03237   0.0248   0.2463   1.0000
  13.000   1.0496   0.04253   0.03384   0.0259   0.2402   1.0000
  13.250   1.0530   0.04423   0.03561   0.0269   0.2341   1.0000
  13.500   1.0630   0.04538   0.03671   0.0277   0.2288   1.0000
  13.750   1.0652   0.04731   0.03873   0.0286   0.2231   1.0000
  14.000   1.0689   0.04914   0.04062   0.0293   0.2176   1.0000
  14.250   1.0800   0.05029   0.04172   0.0300   0.2129   1.0000
  14.500   1.0813   0.05246   0.04399   0.0305   0.2080   1.0000
  14.750   1.0830   0.05464   0.04626   0.0310   0.2030   1.0000
  15.000   1.0912   0.05616   0.04776   0.0315   0.1986   1.0000
  15.250   1.0980   0.05790   0.04952   0.0319   0.1945   1.0000
  15.500   1.0945   0.06079   0.05256   0.0320   0.1902   1.0000
  15.750   1.0975   0.06302   0.05487   0.0322   0.1862   1.0000
  16.000   1.1077   0.06443   0.05626   0.0324   0.1828   1.0000
  16.250   1.1135   0.06641   0.05828   0.0326   0.1794   1.0000
  16.500   1.1022   0.07045   0.06252   0.0321   0.1755   1.0000
  16.750   1.0988   0.07363   0.06580   0.0318   0.1719   1.0000
  17.000   1.1049   0.07564   0.06784   0.0317   0.1687   1.0000
  17.250   1.1254   0.07582   0.06793   0.0322   0.1656   1.0000
  17.500   1.0964   0.08244   0.07484   0.0305   0.1623   1.0000
  17.750   1.0712   0.08887   0.08147   0.0287   0.1587   1.0000
  18.000   1.0642   0.09286   0.08554   0.0275   0.1551   1.0000
  18.250   1.0902   0.09202   0.08459   0.0282   0.1515   1.0000
  18.500   1.0530   0.10057   0.09337   0.0252   0.1483   1.0000
<< Back to RAF 69 AIRFOIL (raf69-il)

Polar data table (+)

Polar graphs


<< Back to RAF 69 AIRFOIL (raf69-il)