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

OAF117 AIRFOIL (oaf117-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: OAF117 AIRFOIL (oaf117-il)
Reynolds number: 200,000
Max Cl/Cd: 64.17 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf117-il-200000.txt
Download as CSV file: xf-oaf117-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF117 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.6110   0.13424   0.13055   0.0305   1.0000   0.0766
 -11.000  -0.8846   0.05990   0.05590  -0.0265   1.0000   0.0471
 -10.750  -0.9196   0.04892   0.04414  -0.0390   1.0000   0.0465
 -10.500  -0.9074   0.04570   0.04084  -0.0402   1.0000   0.0474
 -10.250  -0.8898   0.04352   0.03861  -0.0409   1.0000   0.0483
 -10.000  -0.8755   0.04012   0.03490  -0.0426   1.0000   0.0493
  -9.750  -0.8583   0.03690   0.03129  -0.0441   1.0000   0.0508
  -9.500  -0.8383   0.03385   0.02769  -0.0456   1.0000   0.0525
  -9.250  -0.8162   0.03082   0.02426  -0.0468   1.0000   0.0540
  -9.000  -0.7912   0.02906   0.02248  -0.0472   1.0000   0.0556
  -8.750  -0.7650   0.02772   0.02102  -0.0477   1.0000   0.0577
  -8.500  -0.7379   0.02628   0.01930  -0.0482   1.0000   0.0601
  -8.250  -0.7107   0.02451   0.01724  -0.0488   1.0000   0.0622
  -8.000  -0.6835   0.02313   0.01590  -0.0492   1.0000   0.0646
  -7.750  -0.6553   0.02217   0.01487  -0.0496   1.0000   0.0676
  -7.500  -0.6262   0.02129   0.01374  -0.0500   1.0000   0.0706
  -7.250  -0.5983   0.01983   0.01236  -0.0504   1.0000   0.0739
  -7.000  -0.5693   0.01907   0.01157  -0.0508   1.0000   0.0779
  -6.750  -0.5400   0.01814   0.01053  -0.0512   1.0000   0.0823
  -6.500  -0.5109   0.01733   0.00981  -0.0516   1.0000   0.0872
  -6.250  -0.4812   0.01664   0.00904  -0.0520   1.0000   0.0934
  -6.000  -0.4515   0.01596   0.00844  -0.0526   1.0000   0.1008
  -5.750  -0.4216   0.01528   0.00780  -0.0531   1.0000   0.1099
  -5.500  -0.3914   0.01467   0.00724  -0.0537   1.0000   0.1222
  -5.250  -0.3611   0.01407   0.00671  -0.0543   1.0000   0.1386
  -5.000  -0.3307   0.01360   0.00632  -0.0549   1.0000   0.1585
  -4.750  -0.3002   0.01323   0.00601  -0.0555   1.0000   0.1790
  -4.500  -0.2698   0.01286   0.00573  -0.0562   1.0000   0.1981
  -4.250  -0.2393   0.01255   0.00550  -0.0569   1.0000   0.2165
  -4.000  -0.2087   0.01228   0.00532  -0.0576   1.0000   0.2354
  -3.750  -0.1782   0.01205   0.00516  -0.0584   1.0000   0.2550
  -3.500  -0.1429   0.01188   0.00507  -0.0602   0.9848   0.2765
  -3.250  -0.1022   0.01177   0.00504  -0.0627   0.9538   0.3006
  -3.000  -0.0719   0.01181   0.00506  -0.0627   0.9208   0.3237
  -2.500  -0.0233   0.01187   0.00511  -0.0599   0.8633   0.3647
  -2.250   0.0018   0.01189   0.00508  -0.0587   0.8389   0.3845
  -2.000   0.0286   0.01187   0.00503  -0.0580   0.8146   0.4046
  -1.750   0.0557   0.01186   0.00496  -0.0574   0.7920   0.4243
  -1.500   0.0832   0.01187   0.00490  -0.0568   0.7711   0.4443
  -1.250   0.1114   0.01187   0.00484  -0.0565   0.7511   0.4647
  -1.000   0.1398   0.01184   0.00481  -0.0562   0.7319   0.4842
  -0.750   0.1684   0.01182   0.00478  -0.0560   0.7137   0.5043
  -0.500   0.1972   0.01181   0.00477  -0.0559   0.6968   0.5262
  -0.250   0.2260   0.01182   0.00476  -0.0557   0.6808   0.5514
   0.000   0.2547   0.01179   0.00478  -0.0555   0.6651   0.5763
   0.250   0.2836   0.01177   0.00479  -0.0555   0.6491   0.6045
   0.500   0.3123   0.01172   0.00483  -0.0553   0.6336   0.6360
   0.750   0.3405   0.01167   0.00488  -0.0550   0.6189   0.6729
   1.000   0.3678   0.01158   0.00493  -0.0545   0.6048   0.7183
   1.250   0.3926   0.01142   0.00495  -0.0533   0.5912   0.7848
   1.750   0.4414   0.01100   0.00466  -0.0506   0.5632   1.0000
   2.000   0.4718   0.01117   0.00472  -0.0509   0.5481   1.0000
   2.250   0.5019   0.01135   0.00480  -0.0512   0.5336   1.0000
   2.500   0.5318   0.01154   0.00489  -0.0515   0.5197   1.0000
   2.750   0.5615   0.01175   0.00499  -0.0517   0.5063   1.0000
   3.000   0.5911   0.01195   0.00511  -0.0518   0.4929   1.0000
   3.250   0.6208   0.01213   0.00525  -0.0520   0.4786   1.0000
   3.500   0.6503   0.01233   0.00541  -0.0522   0.4649   1.0000
   3.750   0.6796   0.01255   0.00558  -0.0523   0.4513   1.0000
   4.000   0.7088   0.01278   0.00574  -0.0524   0.4380   1.0000
   4.250   0.7378   0.01303   0.00591  -0.0525   0.4244   1.0000
   4.500   0.7670   0.01323   0.00612  -0.0527   0.4099   1.0000
   4.750   0.7960   0.01346   0.00633  -0.0528   0.3951   1.0000
   5.000   0.8248   0.01370   0.00655  -0.0528   0.3799   1.0000
   5.250   0.8535   0.01394   0.00677  -0.0529   0.3638   1.0000
   5.500   0.8819   0.01420   0.00699  -0.0530   0.3466   1.0000
   5.750   0.9102   0.01448   0.00723  -0.0530   0.3286   1.0000
   6.000   0.9382   0.01480   0.00749  -0.0530   0.3101   1.0000
   6.250   0.9662   0.01513   0.00779  -0.0531   0.2909   1.0000
   6.500   0.9940   0.01549   0.00814  -0.0531   0.2713   1.0000
   6.750   1.0214   0.01592   0.00853  -0.0531   0.2516   1.0000
   7.000   1.0486   0.01638   0.00894  -0.0532   0.2317   1.0000
   7.250   1.0755   0.01688   0.00939  -0.0532   0.2116   1.0000
   7.500   1.1019   0.01745   0.00992  -0.0531   0.1913   1.0000
   7.750   1.1275   0.01815   0.01053  -0.0531   0.1713   1.0000
   8.000   1.1527   0.01892   0.01124  -0.0529   0.1512   1.0000
   8.250   1.1771   0.01977   0.01203  -0.0528   0.1310   1.0000
   8.500   1.1999   0.02086   0.01298  -0.0525   0.1136   1.0000
   8.750   1.2225   0.02193   0.01400  -0.0520   0.0998   1.0000
   9.000   1.2441   0.02307   0.01509  -0.0515   0.0894   1.0000
   9.250   1.2663   0.02405   0.01614  -0.0509   0.0818   1.0000
   9.500   1.2857   0.02538   0.01744  -0.0501   0.0760   1.0000
   9.750   1.3062   0.02643   0.01855  -0.0494   0.0711   1.0000
  10.000   1.3231   0.02798   0.02005  -0.0484   0.0673   1.0000
  10.250   1.3423   0.02911   0.02131  -0.0475   0.0640   1.0000
  10.500   1.3598   0.03036   0.02257  -0.0465   0.0612   1.0000
  10.750   1.3747   0.03217   0.02433  -0.0452   0.0587   1.0000
  11.000   1.3911   0.03341   0.02576  -0.0441   0.0565   1.0000
  11.250   1.4062   0.03480   0.02724  -0.0428   0.0545   1.0000
  11.500   1.4205   0.03626   0.02873  -0.0416   0.0529   1.0000
  11.750   1.4352   0.03821   0.03064  -0.0404   0.0513   1.0000
  12.000   1.4460   0.04011   0.03272  -0.0389   0.0501   1.0000
  12.250   1.4531   0.04184   0.03466  -0.0371   0.0490   1.0000
  12.500   1.4568   0.04366   0.03664  -0.0350   0.0479   1.0000
  12.750   1.4602   0.04566   0.03878  -0.0334   0.0469   1.0000
  13.000   1.4646   0.04775   0.04097  -0.0323   0.0459   1.0000
  13.250   1.4712   0.04992   0.04321  -0.0313   0.0451   1.0000
  13.500   1.4814   0.05245   0.04575  -0.0304   0.0442   1.0000
  13.750   1.4806   0.05586   0.04935  -0.0296   0.0437   1.0000
  14.000   1.4716   0.05937   0.05313  -0.0294   0.0434   1.0000
  14.250   1.4608   0.06340   0.05743  -0.0298   0.0431   1.0000
  14.500   1.4479   0.06797   0.06226  -0.0309   0.0428   1.0000
  14.750   1.4324   0.07314   0.06769  -0.0326   0.0426   1.0000
  15.000   1.4141   0.07897   0.07377  -0.0350   0.0425   1.0000
  15.250   1.3928   0.08554   0.08059  -0.0381   0.0424   1.0000
  15.500   1.3681   0.09298   0.08828  -0.0421   0.0424   1.0000
  15.750   1.3394   0.10152   0.09707  -0.0471   0.0424   1.0000
  16.000   1.3068   0.11145   0.10724  -0.0533   0.0427   1.0000
  16.250   1.2690   0.12343   0.11945  -0.0613   0.0430   1.0000
  16.500   1.2255   0.13789   0.13412  -0.0711   0.0435   1.0000
<< Back to OAF117 AIRFOIL (oaf117-il)

Polar data table (+)

Polar graphs


<< Back to OAF117 AIRFOIL (oaf117-il)