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

DU 86-137/25 AIRFOIL (du861372-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: DU 86-137/25 AIRFOIL (du861372-il)
Reynolds number: 200,000
Max Cl/Cd: 32.77 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-du861372-il-200000.txt
Download as CSV file: xf-du861372-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DU 86-137/25 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
  ------ -------- --------- --------- -------- -------- --------
 -13.500  -0.6830   0.10284   0.09782  -0.0346   0.3565   0.0411
 -13.250  -0.6924   0.09677   0.09172  -0.0378   0.3566   0.0410
 -13.000  -0.7009   0.09133   0.08625  -0.0404   0.3566   0.0407
 -12.750  -0.7255   0.08454   0.07937  -0.0440   0.3568   0.0405
 -12.500  -0.7424   0.07937   0.07412  -0.0459   0.3568   0.0402
 -12.250  -0.7596   0.07464   0.06930  -0.0471   0.3568   0.0396
 -12.000  -0.7853   0.06998   0.06449  -0.0475   0.3570   0.0396
 -11.750  -0.8066   0.06608   0.06043  -0.0468   0.3570   0.0395
 -11.500  -0.8275   0.06268   0.05685  -0.0452   0.3573   0.0396
 -11.250  -0.8537   0.05976   0.05368  -0.0422   0.3573   0.0402
 -11.000  -0.8733   0.05728   0.05099  -0.0384   0.3574   0.0403
 -10.750  -0.8970   0.05525   0.04862  -0.0333   0.3576   0.0408
 -10.500  -0.8695   0.05400   0.04760  -0.0346   0.3578   0.0447
 -10.250  -0.8769   0.05172   0.04503  -0.0314   0.3580   0.0457
 -10.000  -0.8887   0.04980   0.04265  -0.0271   0.3581   0.0468
  -9.750  -0.8640   0.04860   0.04167  -0.0280   0.3583   0.0522
  -9.500  -0.8660   0.04718   0.03982  -0.0245   0.3585   0.0543
  -9.250  -0.8477   0.04588   0.03869  -0.0247   0.3587   0.0622
  -9.000  -0.8408   0.04368   0.03613  -0.0223   0.3589   0.0633
  -8.750  -0.8298   0.04144   0.03354  -0.0202   0.3590   0.0639
  -8.500  -0.8156   0.03938   0.03114  -0.0184   0.3592   0.0644
  -8.250  -0.8005   0.03842   0.02980  -0.0164   0.3594   0.0653
  -8.000  -0.7831   0.03536   0.02641  -0.0153   0.3595   0.0664
  -7.750  -0.7596   0.03308   0.02390  -0.0147   0.3597   0.0666
  -7.500  -0.7362   0.03231   0.02290  -0.0137   0.3601   0.0659
  -7.250  -0.7047   0.02957   0.02006  -0.0143   0.3604   0.0647
  -7.000  -0.6736   0.02750   0.01788  -0.0147   0.3608   0.0639
  -6.750  -0.6421   0.02587   0.01613  -0.0150   0.3611   0.0630
  -6.500  -0.6108   0.02450   0.01468  -0.0154   0.3613   0.0621
  -6.250  -0.5805   0.02329   0.01345  -0.0156   0.3615   0.0610
  -6.000  -0.5539   0.02224   0.01238  -0.0152   0.3616   0.0596
  -5.750  -0.5321   0.02135   0.01145  -0.0141   0.3619   0.0582
  -5.500  -0.5134   0.02059   0.01067  -0.0127   0.3622   0.0578
  -5.250  -0.4960   0.01989   0.00995  -0.0112   0.3624   0.0584
  -5.000  -0.4781   0.01927   0.00927  -0.0097   0.3627   0.0586
  -4.750  -0.4592   0.01871   0.00862  -0.0083   0.3630   0.0587
  -4.500  -0.4388   0.01824   0.00805  -0.0072   0.3633   0.0599
  -4.250  -0.4173   0.01784   0.00756  -0.0062   0.3636   0.0628
  -4.000  -0.3948   0.01752   0.00714  -0.0054   0.3640   0.0674
  -3.750  -0.3724   0.01719   0.00674  -0.0046   0.3644   0.0738
  -3.500  -0.3484   0.01698   0.00644  -0.0041   0.3647   0.0793
  -3.250  -0.3257   0.01659   0.00617  -0.0035   0.3652   0.0957
  -3.000  -0.3006   0.01640   0.00595  -0.0032   0.3659   0.1214
  -2.750  -0.2735   0.01638   0.00578  -0.0031   0.3666   0.0905
  -2.500  -0.2467   0.01635   0.00562  -0.0030   0.3673   0.0861
  -2.250  -0.2201   0.01630   0.00549  -0.0029   0.3682   0.0821
  -2.000  -0.1938   0.01624   0.00536  -0.0028   0.3690   0.0787
  -1.750  -0.1676   0.01616   0.00525  -0.0026   0.3700   0.0775
  -1.500  -0.1421   0.01602   0.00512  -0.0024   0.3710   0.0847
  -1.250  -0.1153   0.01601   0.00503  -0.0024   0.3721   0.0779
  -1.000  -0.0889   0.01596   0.00495  -0.0023   0.3730   0.0761
  -0.750  -0.0626   0.01590   0.00486  -0.0022   0.3738   0.0754
  -0.500  -0.0363   0.01584   0.00478  -0.0021   0.3745   0.0763
  -0.250  -0.0111   0.01567   0.00471  -0.0019   0.3753   0.0946
   0.000   0.0147   0.01552   0.00464  -0.0018   0.3754   0.1353
   0.250   0.0360   0.01483   0.00441  -0.0012   0.3749   0.2567
   0.500   0.0412   0.01309   0.00455   0.0025   0.3746   0.6731
   0.750   0.0678   0.01313   0.00467   0.0027   0.3742   0.6946
   1.000   0.0945   0.01321   0.00477   0.0029   0.3747   0.7107
   1.250   0.1210   0.01331   0.00491   0.0032   0.3752   0.7269
   1.500   0.1473   0.01345   0.00509   0.0035   0.3757   0.7430
   1.750   0.1731   0.01362   0.00533   0.0041   0.3759   0.7594
   2.000   0.1973   0.01393   0.00575   0.0053   0.3756   0.7803
   2.250   0.2219   0.01414   0.00606   0.0063   0.3749   0.7930
   2.500   0.2446   0.01443   0.00646   0.0080   0.3743   0.8086
   2.750   0.2645   0.01479   0.00698   0.0106   0.3730   0.8266
   3.000   0.2868   0.01495   0.00723   0.0122   0.3717   0.8390
   3.250   0.3047   0.01520   0.00762   0.0153   0.3690   0.8554
   3.500   0.3185   0.01545   0.00804   0.0200   0.3632   0.8755
   3.750   0.3387   0.01543   0.00810   0.0220   0.3268   0.8880
   4.000   0.3547   0.01584   0.00815   0.0238   0.3195   0.8915
   4.250   0.3734   0.01659   0.00866   0.0249   0.3138   0.8950
   4.500   0.4082   0.01822   0.00977   0.0230   0.3051   0.8973
   4.750   0.4401   0.01953   0.01099   0.0216   0.2955   0.9004
   5.000   0.4658   0.02056   0.01209   0.0213   0.2837   0.9036
   5.250   0.4905   0.02178   0.01342   0.0213   0.2694   0.9061
   5.500   0.5106   0.02206   0.01393   0.0223   0.2566   0.9098
   5.750   0.5301   0.02217   0.01426   0.0234   0.2454   0.9140
   6.000   0.5485   0.02217   0.01443   0.0245   0.2364   0.9182
   6.250   0.5630   0.02094   0.01359   0.0269   0.2213   0.9222
   6.500   0.5815   0.01966   0.01263   0.0287   0.1880   0.9269
   6.750   0.5921   0.01864   0.01184   0.0314   0.1724   0.9332
   7.000   0.6086   0.01862   0.01182   0.0329   0.1462   0.9378
   7.250   0.6240   0.01907   0.01219   0.0343   0.1240   0.9427
   7.500   0.6394   0.01951   0.01257   0.0357   0.1061   0.9486
   7.750   0.6522   0.02014   0.01314   0.0373   0.0934   0.9547
   8.000   0.6684   0.02091   0.01390   0.0380   0.0822   0.9608
   8.250   0.6861   0.02176   0.01481   0.0385   0.0736   0.9675
   8.500   0.7048   0.02334   0.01632   0.0378   0.0660   0.9737
   8.750   0.7345   0.02410   0.01720   0.0364   0.0597   0.9790
   9.000   0.7596   0.02541   0.01845   0.0350   0.0547   0.9848
   9.250   0.7888   0.02685   0.01999   0.0339   0.0508   0.9900
   9.500   0.8200   0.02811   0.02135   0.0326   0.0479   0.9974
   9.750   0.8351   0.02899   0.02227   0.0338   0.0457   1.0000
  10.000   0.8551   0.03039   0.02362   0.0341   0.0431   1.0000
  10.250   0.8868   0.03278   0.02622   0.0329   0.0409   1.0000
  10.500   0.9100   0.03470   0.02840   0.0328   0.0397   1.0000
  10.750   0.9290   0.03695   0.03092   0.0331   0.0386   1.0000
  11.000   0.9420   0.03948   0.03374   0.0338   0.0378   1.0000
  11.250   0.9493   0.04198   0.03651   0.0350   0.0370   1.0000
  11.500   0.9565   0.04378   0.03843   0.0361   0.0358   1.0000
  11.750   0.9625   0.04550   0.04024   0.0370   0.0347   1.0000
  12.000   0.9633   0.04791   0.04279   0.0381   0.0340   1.0000
  12.250   0.9560   0.05118   0.04629   0.0394   0.0338   1.0000
  12.500   0.9438   0.05471   0.05006   0.0405   0.0338   1.0000
  12.750   0.9232   0.05897   0.05462   0.0412   0.0340   1.0000
  13.000   0.8588   0.06819   0.06439   0.0404   0.0356   1.0000
  13.250   0.7903   0.08085   0.07741   0.0344   0.0372   1.0000
  13.500   0.7237   0.09867   0.09535   0.0232   0.0404   1.0000
  13.750   0.7023   0.10840   0.10505   0.0182   0.0417   1.0000
<< Back to DU 86-137/25 AIRFOIL (du861372-il)

Polar data table (+)

Polar graphs


<< Back to DU 86-137/25 AIRFOIL (du861372-il)