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

RAF 6 AIRFOIL (raf6-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RAF 6 AIRFOIL (raf6-il)
Reynolds number: 200,000
Max Cl/Cd: 70.26 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf6-il-200000.txt
Download as CSV file: xf-raf6-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 6 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3372   0.08516   0.08226  -0.0216   1.0000   0.0352
  -7.500  -0.3486   0.08314   0.08030  -0.0195   1.0000   0.0355
  -7.250  -0.3629   0.08119   0.07840  -0.0175   1.0000   0.0358
  -7.000  -0.3725   0.07829   0.07553  -0.0175   0.9984   0.0360
  -6.750  -0.4483   0.08152   0.07845  -0.0249   1.0000   0.0343
  -6.500  -0.4490   0.07909   0.07605  -0.0225   1.0000   0.0348
  -6.250  -0.4338   0.07549   0.07242  -0.0246   0.9977   0.0357
  -6.000  -0.4086   0.07102   0.06788  -0.0299   0.9942   0.0370
  -5.750  -0.3833   0.06649   0.06324  -0.0350   0.9895   0.0387
  -5.500  -0.3337   0.06208   0.05823  -0.0451   0.9847   0.0423
  -5.250  -0.3152   0.05636   0.05231  -0.0475   0.9795   0.0430
  -5.000  -0.2953   0.05210   0.04814  -0.0491   0.9765   0.0441
  -4.750  -0.2677   0.04902   0.04500  -0.0516   0.9740   0.0459
  -4.500  -0.2433   0.04624   0.04207  -0.0529   0.9691   0.0485
  -4.250  -0.2004   0.04644   0.04144  -0.0547   0.9639   0.0538
  -4.000  -0.1773   0.03952   0.03472  -0.0578   0.9620   0.0567
  -3.750  -0.1473   0.03743   0.03252  -0.0594   0.9574   0.0621
  -3.500  -0.1126   0.03470   0.02932  -0.0611   0.9519   0.0697
  -3.250  -0.0736   0.03232   0.02687  -0.0641   0.9487   0.0756
  -3.000  -0.0408   0.03018   0.02443  -0.0655   0.9434   0.0863
  -2.750  -0.0057   0.02833   0.02237  -0.0673   0.9379   0.1013
  -2.000   0.1319   0.02136   0.01385  -0.0716   0.9227   0.0544
  -1.750   0.1823   0.01899   0.01137  -0.0754   0.9197   0.0568
  -1.500   0.2111   0.01793   0.01026  -0.0750   0.9106   0.0580
  -1.250   0.2532   0.01675   0.00909  -0.0772   0.9075   0.0609
  -1.000   0.2855   0.01609   0.00840  -0.0776   0.9007   0.0662
  -0.750   0.3227   0.01489   0.00728  -0.0791   0.8956   0.0715
  -0.500   0.3656   0.01398   0.00639  -0.0817   0.8921   0.0836
  -0.250   0.3941   0.01337   0.00582  -0.0814   0.8825   0.1048
   0.000   0.4586   0.01063   0.00555  -0.0880   0.8797   1.0000
   0.250   0.4924   0.01036   0.00515  -0.0887   0.8680   1.0000
   0.500   0.5286   0.01011   0.00477  -0.0900   0.8555   1.0000
   0.750   0.5644   0.00991   0.00446  -0.0912   0.8407   1.0000
   1.000   0.5990   0.00978   0.00419  -0.0921   0.8224   1.0000
   1.250   0.6303   0.00974   0.00401  -0.0923   0.8009   1.0000
   1.500   0.6554   0.00978   0.00394  -0.0913   0.7753   1.0000
   1.750   0.6782   0.00985   0.00388  -0.0899   0.7448   1.0000
   2.000   0.6991   0.00995   0.00381  -0.0880   0.6999   1.0000
   2.250   0.7177   0.01023   0.00369  -0.0855   0.6233   1.0000
   2.500   0.7300   0.01088   0.00377  -0.0820   0.5323   1.0000
   2.750   0.7432   0.01159   0.00402  -0.0790   0.4674   1.0000
   3.000   0.7596   0.01222   0.00431  -0.0768   0.4237   1.0000
   3.250   0.7784   0.01276   0.00462  -0.0750   0.3902   1.0000
   3.500   0.7985   0.01326   0.00492  -0.0736   0.3641   1.0000
   3.750   0.8198   0.01370   0.00521  -0.0724   0.3437   1.0000
   4.000   0.8419   0.01412   0.00551  -0.0713   0.3278   1.0000
   4.250   0.8642   0.01456   0.00583  -0.0704   0.3152   1.0000
   4.500   0.8873   0.01495   0.00616  -0.0695   0.3040   1.0000
   4.750   0.9108   0.01534   0.00650  -0.0688   0.2942   1.0000
   5.000   0.9341   0.01584   0.00688  -0.0681   0.2863   1.0000
   5.250   0.9581   0.01618   0.00725  -0.0674   0.2782   1.0000
   5.500   0.9821   0.01671   0.00768  -0.0669   0.2713   1.0000
   5.750   1.0061   0.01706   0.00809  -0.0663   0.2644   1.0000
   6.000   1.0305   0.01756   0.00852  -0.0658   0.2584   1.0000
   6.250   1.0547   0.01801   0.00903  -0.0653   0.2524   1.0000
   6.500   1.0786   0.01844   0.00946  -0.0647   0.2463   1.0000
   6.750   1.1040   0.01907   0.01003  -0.0645   0.2410   1.0000
   7.000   1.1275   0.01947   0.01056  -0.0638   0.2358   1.0000
   7.250   1.1515   0.01997   0.01107  -0.0634   0.2306   1.0000
   7.500   1.1767   0.02064   0.01173  -0.0632   0.2255   1.0000
   7.750   1.1991   0.02106   0.01229  -0.0623   0.2202   1.0000
   8.000   1.2230   0.02160   0.01286  -0.0619   0.2155   1.0000
   8.250   1.2488   0.02240   0.01366  -0.0619   0.2110   1.0000
   8.500   1.2703   0.02289   0.01434  -0.0609   0.2065   1.0000
   8.750   1.2928   0.02340   0.01491  -0.0601   0.2018   1.0000
   9.000   1.3154   0.02405   0.01557  -0.0596   0.1962   1.0000
   9.250   1.3312   0.02426   0.01597  -0.0576   0.1900   1.0000
   9.500   1.3502   0.02460   0.01627  -0.0564   0.1841   1.0000
   9.750   1.3662   0.02497   0.01684  -0.0545   0.1785   1.0000
  10.000   1.3825   0.02526   0.01724  -0.0527   0.1732   1.0000
  10.250   1.4007   0.02578   0.01775  -0.0515   0.1683   1.0000
  10.500   1.4131   0.02606   0.01829  -0.0490   0.1632   1.0000
  10.750   1.4232   0.02625   0.01855  -0.0461   0.1578   1.0000
  11.000   1.4348   0.02666   0.01909  -0.0437   0.1524   1.0000
  11.250   1.4451   0.02694   0.01955  -0.0411   0.1460   1.0000
  11.500   1.4552   0.02733   0.02006  -0.0387   0.1389   1.0000
  11.750   1.4649   0.02767   0.02051  -0.0363   0.1279   1.0000
  12.000   1.4764   0.02823   0.02113  -0.0344   0.1058   1.0000
  12.250   1.4758   0.03011   0.02274  -0.0314   0.0640   1.0000
  12.500   1.4658   0.03291   0.02531  -0.0277   0.0422   1.0000
  12.750   1.4605   0.03535   0.02781  -0.0247   0.0368   1.0000
  13.000   1.4550   0.03788   0.03049  -0.0222   0.0340   1.0000
  13.250   1.4506   0.04043   0.03324  -0.0201   0.0323   1.0000
  13.500   1.4432   0.04340   0.03639  -0.0184   0.0310   1.0000
  13.750   1.4328   0.04686   0.04002  -0.0170   0.0300   1.0000
  14.000   1.4196   0.05088   0.04420  -0.0163   0.0293   1.0000
  14.250   1.4037   0.05555   0.04903  -0.0163   0.0287   1.0000
  14.500   1.3865   0.06082   0.05446  -0.0170   0.0282   1.0000
  14.750   1.3691   0.06657   0.06035  -0.0185   0.0279   1.0000
  15.000   1.3567   0.07201   0.06597  -0.0203   0.0275   1.0000
  15.250   1.3433   0.07785   0.07198  -0.0224   0.0272   1.0000
  15.500   1.3300   0.08381   0.07809  -0.0247   0.0269   1.0000
  15.750   1.3171   0.08970   0.08413  -0.0270   0.0265   1.0000
<< Back to RAF 6 AIRFOIL (raf6-il)

Polar data table (+)

Polar graphs


<< Back to RAF 6 AIRFOIL (raf6-il)