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

RG 15A 2.5/13.0 AIRFOIL (rg15a213-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RG 15A 2.5/13.0 AIRFOIL (rg15a213-il)
Reynolds number: 100,000
Max Cl/Cd: 50.62 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rg15a213-il-100000.txt
Download as CSV file: xf-rg15a213-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 15A 2.5/13.0 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4013   0.09750   0.09252  -0.0366   1.0000   0.1429
  -9.250  -0.3265   0.08846   0.08392  -0.0362   1.0000   0.1559
  -9.000  -0.5693   0.06841   0.06359  -0.0544   1.0000   0.0785
  -8.750  -0.5793   0.06499   0.06019  -0.0524   1.0000   0.0769
  -8.500  -0.6579   0.05637   0.05070  -0.0491   1.0000   0.0677
  -8.250  -0.6587   0.05224   0.04650  -0.0473   1.0000   0.0670
  -8.000  -0.6603   0.04851   0.04258  -0.0452   1.0000   0.0664
  -7.750  -0.6601   0.04485   0.03862  -0.0432   1.0000   0.0658
  -7.500  -0.6567   0.04130   0.03468  -0.0413   1.0000   0.0655
  -7.250  -0.6488   0.03813   0.03106  -0.0396   1.0000   0.0656
  -7.000  -0.6375   0.03559   0.02799  -0.0380   1.0000   0.0670
  -6.750  -0.6234   0.03298   0.02508  -0.0368   1.0000   0.0691
  -6.500  -0.6063   0.03138   0.02339  -0.0357   1.0000   0.0717
  -6.250  -0.5878   0.02975   0.02154  -0.0345   1.0000   0.0742
  -6.000  -0.5683   0.02820   0.01966  -0.0333   1.0000   0.0775
  -5.750  -0.5489   0.02672   0.01804  -0.0323   1.0000   0.0825
  -5.500  -0.5297   0.02577   0.01704  -0.0313   1.0000   0.0895
  -5.250  -0.5099   0.02446   0.01571  -0.0302   1.0000   0.0980
  -5.000  -0.4905   0.02346   0.01474  -0.0292   1.0000   0.1120
  -4.750  -0.4714   0.02271   0.01401  -0.0283   1.0000   0.1329
  -4.500  -0.4522   0.02206   0.01342  -0.0274   1.0000   0.1541
  -4.250  -0.4107   0.02183   0.01333  -0.0310   0.9934   0.1881
  -4.000  -0.3703   0.02149   0.01310  -0.0341   0.9857   0.2263
  -3.750  -0.3321   0.02091   0.01280  -0.0367   0.9781   0.2698
  -3.500  -0.2914   0.02031   0.01261  -0.0397   0.9713   0.3404
  -3.250  -0.2593   0.01958   0.01246  -0.0410   0.9621   0.4376
  -3.000  -0.2275   0.01929   0.01290  -0.0414   0.9537   0.5692
  -2.750  -0.1929   0.01964   0.01343  -0.0418   0.9449   0.6645
  -2.500  -0.1653   0.01995   0.01375  -0.0409   0.9342   0.7166
  -2.250  -0.1343   0.02027   0.01404  -0.0404   0.9250   0.7575
  -2.000  -0.1041   0.02050   0.01424  -0.0396   0.9157   0.7913
  -1.750  -0.0831   0.02066   0.01438  -0.0371   0.9045   0.8223
  -1.500  -0.0537   0.02074   0.01442  -0.0359   0.8965   0.8527
  -1.250  -0.0296   0.02073   0.01436  -0.0340   0.8859   0.8782
  -1.000  -0.0013   0.02071   0.01427  -0.0331   0.8757   0.9012
  -0.750   0.0477   0.02061   0.01408  -0.0358   0.8695   0.9230
  -0.500   0.0990   0.02061   0.01398  -0.0396   0.8612   0.9429
  -0.250   0.1771   0.02045   0.01370  -0.0482   0.8570   0.9568
   0.000   0.2766   0.02002   0.01315  -0.0611   0.8551   0.9646
   0.250   0.3617   0.01933   0.01238  -0.0714   0.8520   0.9730
   0.500   0.4216   0.01888   0.01189  -0.0775   0.8408   0.9831
   0.750   0.4897   0.01817   0.01114  -0.0850   0.8317   0.9905
   1.000   0.5516   0.01745   0.01037  -0.0913   0.8211   0.9976
   1.250   0.5837   0.01714   0.01002  -0.0923   0.8061   1.0000
   1.500   0.6060   0.01694   0.00978  -0.0913   0.7905   1.0000
   1.750   0.6273   0.01678   0.00959  -0.0900   0.7748   1.0000
   2.000   0.6474   0.01666   0.00942  -0.0886   0.7590   1.0000
   2.250   0.6665   0.01659   0.00930  -0.0870   0.7431   1.0000
   2.500   0.6846   0.01655   0.00922  -0.0852   0.7271   1.0000
   2.750   0.7017   0.01656   0.00918  -0.0832   0.7111   1.0000
   3.000   0.7178   0.01659   0.00916  -0.0810   0.6952   1.0000
   3.250   0.7330   0.01665   0.00916  -0.0787   0.6792   1.0000
   3.500   0.7480   0.01674   0.00921  -0.0763   0.6633   1.0000
   3.750   0.7641   0.01686   0.00927  -0.0741   0.6472   1.0000
   4.000   0.7818   0.01703   0.00937  -0.0722   0.6306   1.0000
   4.250   0.8017   0.01723   0.00953  -0.0707   0.6132   1.0000
   4.500   0.8235   0.01745   0.00971  -0.0695   0.5954   1.0000
   4.750   0.8464   0.01768   0.00989  -0.0685   0.5776   1.0000
   5.000   0.8698   0.01792   0.01008  -0.0675   0.5595   1.0000
   5.250   0.8937   0.01818   0.01029  -0.0667   0.5415   1.0000
   5.500   0.9179   0.01847   0.01050  -0.0659   0.5235   1.0000
   5.750   0.9423   0.01878   0.01072  -0.0651   0.5056   1.0000
   6.000   0.9632   0.01915   0.01109  -0.0638   0.4864   1.0000
   6.250   0.9847   0.01952   0.01146  -0.0626   0.4673   1.0000
   6.500   1.0065   0.01991   0.01180  -0.0614   0.4484   1.0000
   6.750   1.0285   0.02032   0.01213  -0.0603   0.4296   1.0000
   7.000   1.0501   0.02078   0.01252  -0.0592   0.4110   1.0000
   7.250   1.0687   0.02126   0.01303  -0.0576   0.3916   1.0000
   7.500   1.0879   0.02177   0.01353  -0.0561   0.3725   1.0000
   7.750   1.1071   0.02232   0.01403  -0.0547   0.3540   1.0000
   8.000   1.1263   0.02291   0.01457  -0.0533   0.3360   1.0000
   8.250   1.1454   0.02358   0.01515  -0.0519   0.3184   1.0000
   8.500   1.1619   0.02427   0.01589  -0.0502   0.3007   1.0000
   8.750   1.1779   0.02501   0.01663  -0.0484   0.2833   1.0000
   9.000   1.1935   0.02580   0.01743  -0.0466   0.2665   1.0000
   9.250   1.2086   0.02664   0.01826  -0.0447   0.2504   1.0000
   9.500   1.2218   0.02747   0.01907  -0.0426   0.2346   1.0000
   9.750   1.2332   0.02831   0.01991  -0.0403   0.2194   1.0000
  10.000   1.2437   0.02922   0.02080  -0.0380   0.2052   1.0000
  10.250   1.2532   0.03017   0.02177  -0.0356   0.1923   1.0000
  10.500   1.2633   0.03123   0.02279  -0.0333   0.1804   1.0000
  10.750   1.2754   0.03238   0.02384  -0.0314   0.1688   1.0000
  11.000   1.2794   0.03354   0.02517  -0.0284   0.1585   1.0000
  11.250   1.2871   0.03491   0.02660  -0.0261   0.1485   1.0000
  11.500   1.2978   0.03632   0.02792  -0.0244   0.1389   1.0000
  11.750   1.3021   0.03778   0.02954  -0.0219   0.1303   1.0000
  12.000   1.3109   0.03959   0.03142  -0.0202   0.1223   1.0000
  12.250   1.3224   0.04117   0.03292  -0.0189   0.1144   1.0000
  12.500   1.3247   0.04327   0.03526  -0.0167   0.1084   1.0000
  12.750   1.3339   0.04499   0.03700  -0.0153   0.1027   1.0000
  13.000   1.3457   0.04744   0.03955  -0.0143   0.0978   1.0000
  13.250   1.3445   0.04988   0.04228  -0.0123   0.0944   1.0000
  13.500   1.3498   0.05205   0.04458  -0.0111   0.0909   1.0000
  13.750   1.3687   0.05470   0.04717  -0.0110   0.0868   1.0000
  14.000   1.3550   0.05770   0.05054  -0.0088   0.0854   1.0000
  14.250   1.3419   0.06112   0.05427  -0.0074   0.0838   1.0000
  14.500   1.3287   0.06490   0.05834  -0.0065   0.0825   1.0000
  14.750   1.3141   0.06905   0.06276  -0.0061   0.0814   1.0000
  15.000   1.2978   0.07360   0.06754  -0.0063   0.0805   1.0000
  15.250   1.2765   0.07891   0.07311  -0.0073   0.0799   1.0000
  15.500   1.2398   0.08642   0.08093  -0.0098   0.0803   1.0000
  15.750   1.1773   0.09863   0.09351  -0.0162   0.0824   1.0000
  16.000   1.1047   0.11541   0.11052  -0.0265   0.0852   1.0000
  16.250   1.0373   0.13516   0.13037  -0.0385   0.0875   1.0000
<< Back to RG 15A 2.5/13.0 AIRFOIL (rg15a213-il)

Polar data table (+)

Polar graphs


<< Back to RG 15A 2.5/13.0 AIRFOIL (rg15a213-il)