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

FX 62-K-153/20 AIRFOIL (fx62k153-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: FX 62-K-153/20 AIRFOIL (fx62k153-il)
Reynolds number: 100,000
Max Cl/Cd: 45.26 at α=9.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-fx62k153-il-100000.txt
Download as CSV file: xf-fx62k153-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 62-K-153/20 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3151   0.10664   0.10286  -0.0498   0.9678   0.1156
  -8.250  -0.3430   0.10334   0.09964  -0.0512   0.9670   0.1186
  -8.000  -0.4086   0.10235   0.09834  -0.0541   0.9743   0.1115
  -7.750  -0.4223   0.09965   0.09569  -0.0537   0.9712   0.1146
  -7.500  -0.4491   0.09687   0.09299  -0.0532   0.9686   0.1163
  -7.250  -0.4792   0.09305   0.08921  -0.0551   0.9658   0.1172
  -7.000  -0.5154   0.08841   0.08445  -0.0605   0.9649   0.1188
  -6.750  -0.5148   0.08356   0.07961  -0.0618   0.9629   0.1225
  -6.500  -0.5000   0.08078   0.07684  -0.0620   0.9607   0.1297
  -6.250  -0.5124   0.07596   0.07181  -0.0658   0.9609   0.1373
  -6.000  -0.5070   0.07368   0.06959  -0.0635   0.9588   0.1432
  -5.750  -0.6092   0.07675   0.07295  -0.0389   1.0001   0.1261
  -5.500  -0.6018   0.07166   0.06761  -0.0438   1.0001   0.1378
  -5.250  -0.5886   0.06778   0.06353  -0.0469   1.0001   0.1521
  -5.000  -0.5741   0.06446   0.06015  -0.0483   1.0001   0.1676
  -4.750  -0.3249   0.04319   0.03541  -0.0906   0.9444   0.0674
  -4.500  -0.2942   0.04130   0.03314  -0.0918   0.9417   0.0693
  -4.250  -0.2557   0.03943   0.03086  -0.0937   0.9378   0.0700
  -4.000  -0.2040   0.03857   0.02955  -0.0977   0.9325   0.0759
  -3.750  -0.1855   0.03719   0.02822  -0.0964   0.9290   0.0787
  -3.500  -0.1585   0.03646   0.02755  -0.0966   0.9256   0.0843
  -3.250  -0.1179   0.03576   0.02694  -0.0993   0.9213   0.0976
  -3.000  -0.0752   0.03506   0.02640  -0.1026   0.9165   0.1240
  -2.750  -0.0422   0.03430   0.02831  -0.1032   0.9123   0.6540
  -2.500  -0.0413   0.03605   0.02999  -0.0958   0.9082   0.6936
  -2.250  -0.0335   0.03769   0.03155  -0.0891   0.9037   0.7241
  -2.000  -0.0412   0.03841   0.03226  -0.0818   0.8989   0.7428
  -1.750  -0.0398   0.03924   0.03303  -0.0752   0.8944   0.7670
  -1.500  -0.0303   0.04013   0.03385  -0.0688   0.8898   0.7964
  -1.250  -0.0429   0.04021   0.03394  -0.0611   0.8844   0.8145
  -1.000  -0.0419   0.04039   0.03407  -0.0548   0.8796   0.8396
  -0.750  -0.0309   0.04049   0.03405  -0.0504   0.8733   0.8615
  -0.500  -0.0192   0.04048   0.03395  -0.0485   0.8670   0.8711
  -0.250   0.0275   0.04089   0.03413  -0.0525   0.8598   0.8772
   0.000   0.0409   0.04098   0.03414  -0.0516   0.8522   0.8823
   0.250   0.0879   0.04130   0.03427  -0.0552   0.8461   0.8872
   0.500   0.1010   0.04153   0.03444  -0.0546   0.8370   0.8919
   0.750   0.1497   0.04186   0.03461  -0.0587   0.8315   0.8960
   1.000   0.1581   0.04204   0.03476  -0.0569   0.8218   0.9004
   1.250   0.2068   0.04236   0.03495  -0.0609   0.8167   0.9046
   1.500   0.2215   0.04282   0.03539  -0.0606   0.8066   0.9084
   1.750   0.2635   0.04296   0.03544  -0.0632   0.8020   0.9123
   2.000   0.2728   0.04346   0.03594  -0.0619   0.7920   0.9163
   2.250   0.3182   0.04375   0.03616  -0.0654   0.7876   0.9199
   2.500   0.3315   0.04435   0.03677  -0.0648   0.7773   0.9237
   2.750   0.3725   0.04450   0.03687  -0.0673   0.7729   0.9273
   3.000   0.3838   0.04528   0.03768  -0.0665   0.7629   0.9312
   3.250   0.4271   0.04549   0.03787  -0.0696   0.7585   0.9350
   3.500   0.4395   0.04619   0.03860  -0.0688   0.7480   0.9389
   3.750   0.4804   0.04626   0.03868  -0.0712   0.7438   0.9427
   4.000   0.4945   0.04705   0.03952  -0.0707   0.7330   0.9471
   4.250   0.5401   0.04682   0.03930  -0.0734   0.7289   0.9513
   4.500   0.5543   0.04743   0.03997  -0.0728   0.7173   0.9558
   4.750   0.6019   0.04703   0.03962  -0.0755   0.7139   0.9606
   5.000   0.6144   0.04774   0.04040  -0.0748   0.7020   0.9662
   5.250   0.6334   0.04838   0.04111  -0.0748   0.6916   0.9722
   5.500   0.6773   0.04785   0.04067  -0.0771   0.6869   0.9786
   5.750   0.6957   0.04843   0.04134  -0.0770   0.6752   0.9912
   6.000   0.7486   0.04741   0.04042  -0.0802   0.6716   0.9999
   6.250   0.7700   0.04818   0.04131  -0.0808   0.6595   0.9999
   6.500   0.8242   0.04704   0.04030  -0.0840   0.6569   0.9999
   6.750   0.8430   0.04801   0.04139  -0.0844   0.6445   0.9999
   7.000   0.8662   0.04875   0.04227  -0.0851   0.6332   0.9999
   7.250   0.9197   0.04713   0.04083  -0.0878   0.6295   0.9999
   7.500   0.9477   0.04710   0.04094  -0.0883   0.6176   0.9999
   7.750   1.0175   0.04297   0.03705  -0.0910   0.6149   0.9999
   8.000   1.0896   0.03804   0.03236  -0.0934   0.6135   0.9999
   8.250   1.1258   0.03642   0.03096  -0.0936   0.6008   0.9999
   8.500   1.1710   0.03400   0.02875  -0.0944   0.5883   0.9999
   8.750   1.2225   0.03121   0.02616  -0.0957   0.5732   0.9999
   9.000   1.2621   0.02965   0.02477  -0.0962   0.5504   0.9999
   9.250   1.2912   0.02902   0.02421  -0.0958   0.5188   0.9999
   9.500   1.3124   0.02900   0.02412  -0.0947   0.4761   0.9999
   9.750   1.3215   0.02999   0.02490  -0.0925   0.4251   0.9999
  10.000   1.3210   0.03182   0.02643  -0.0898   0.3680   0.9999
  10.250   1.3135   0.03435   0.02856  -0.0867   0.3098   0.9999
  10.500   1.3039   0.03727   0.03105  -0.0838   0.2596   0.9999
  10.750   1.2962   0.04027   0.03373  -0.0813   0.2174   0.9999
  11.000   1.2923   0.04313   0.03633  -0.0792   0.1844   0.9999
  11.250   1.2915   0.04583   0.03878  -0.0775   0.1602   0.9999
  11.500   1.2958   0.04822   0.04104  -0.0761   0.1399   0.9999
  11.750   1.3040   0.05038   0.04310  -0.0748   0.1240   0.9999
  12.000   1.3155   0.05237   0.04503  -0.0737   0.1107   0.9999
  12.250   1.3302   0.05422   0.04684  -0.0728   0.0996   0.9999
  12.500   1.3563   0.05570   0.04819  -0.0723   0.0896   0.9999
  12.750   1.3679   0.05780   0.05047  -0.0715   0.0821   0.9999
  13.000   1.3771   0.06015   0.05294  -0.0708   0.0744   0.9999
  13.250   1.3789   0.06254   0.05530  -0.0701   0.0674   0.9999
  13.500   1.3750   0.06556   0.05860  -0.0691   0.0613   0.9999
  13.750   1.3803   0.06816   0.06106  -0.0685   0.0540   0.9999
  14.000   1.3765   0.07157   0.06489  -0.0674   0.0493   0.9999
  14.250   1.3861   0.07467   0.06790  -0.0664   0.0428   0.9999
  14.500   1.3796   0.07861   0.07228  -0.0654   0.0397   0.9999
  14.750   1.3756   0.08220   0.07607  -0.0648   0.0368   0.9999
  15.000   1.3832   0.08570   0.07947  -0.0643   0.0335   0.9999
  15.250   1.3699   0.09016   0.08432  -0.0641   0.0327   0.9999
  15.500   1.3572   0.09505   0.08957  -0.0643   0.0320   0.9999
  15.750   1.3433   0.10033   0.09519  -0.0649   0.0314   0.9999
  16.000   1.3275   0.10609   0.10127  -0.0662   0.0308   0.9999
  16.250   1.3106   0.11232   0.10780  -0.0681   0.0307   0.9999
  16.500   1.2910   0.11928   0.11505  -0.0710   0.0305   0.9999
  16.750   1.2713   0.12669   0.12272  -0.0746   0.0308   0.9999
  17.000   1.2502   0.13487   0.13116  -0.0793   0.0311   0.9999
  17.250   1.2277   0.14400   0.14051  -0.0852   0.0315   0.9999
  17.500   1.2066   0.15361   0.15029  -0.0917   0.0322   0.9999
  17.750   1.1868   0.16365   0.16045  -0.0987   0.0329   0.9999
  18.000   1.1688   0.17412   0.17100  -0.1060   0.0334   0.9999
<< Back to FX 62-K-153/20 AIRFOIL (fx62k153-il)

Polar data table (+)

Polar graphs


<< Back to FX 62-K-153/20 AIRFOIL (fx62k153-il)