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

CLARK Z AIRFOIL (clarkz-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: CLARK Z AIRFOIL (clarkz-il)
Reynolds number: 200,000
Max Cl/Cd: 72.15 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-clarkz-il-200000-n5.txt
Download as CSV file: xf-clarkz-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CLARK Z 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3309   0.09910   0.09533  -0.0418   1.0000   0.0457
  -9.250  -0.3279   0.09714   0.09342  -0.0407   1.0000   0.0464
  -8.750  -0.3533   0.08562   0.08195  -0.0467   0.9955   0.0378
  -8.500  -0.3388   0.08192   0.07825  -0.0503   0.9903   0.0374
  -8.250  -0.3261   0.07713   0.07346  -0.0561   0.9840   0.0368
  -8.000  -0.3151   0.07120   0.06751  -0.0643   0.9742   0.0364
  -7.750  -0.3038   0.06382   0.06005  -0.0744   0.9638   0.0364
  -7.500  -0.2924   0.05565   0.05170  -0.0838   0.9537   0.0364
  -7.250  -0.2806   0.04754   0.04330  -0.0906   0.9447   0.0360
  -7.000  -0.2747   0.03923   0.03448  -0.0941   0.9332   0.0357
  -6.750  -0.2616   0.03289   0.02752  -0.0958   0.9253   0.0357
  -6.500  -0.2446   0.02913   0.02326  -0.0958   0.9165   0.0359
  -6.250  -0.2213   0.02641   0.02005  -0.0962   0.9104   0.0367
  -6.000  -0.1999   0.02442   0.01763  -0.0956   0.9017   0.0373
  -5.750  -0.1731   0.02271   0.01555  -0.0957   0.8960   0.0377
  -5.500  -0.1493   0.02145   0.01400  -0.0951   0.8873   0.0379
  -5.250  -0.1219   0.02031   0.01259  -0.0951   0.8808   0.0383
  -5.000  -0.0966   0.01929   0.01147  -0.0947   0.8725   0.0387
  -4.750  -0.0693   0.01847   0.01055  -0.0946   0.8652   0.0392
  -4.500  -0.0430   0.01778   0.00978  -0.0942   0.8569   0.0397
  -4.250  -0.0157   0.01714   0.00904  -0.0940   0.8489   0.0403
  -4.000   0.0108   0.01660   0.00843  -0.0936   0.8400   0.0412
  -3.750   0.0382   0.01609   0.00782  -0.0933   0.8316   0.0425
  -3.500   0.0646   0.01562   0.00727  -0.0929   0.8220   0.0435
  -3.250   0.0921   0.01515   0.00669  -0.0925   0.8137   0.0444
  -3.000   0.1178   0.01466   0.00619  -0.0920   0.8035   0.0452
  -2.750   0.1446   0.01425   0.00577  -0.0917   0.7947   0.0463
  -2.500   0.1710   0.01393   0.00542  -0.0913   0.7847   0.0477
  -2.250   0.1977   0.01364   0.00509  -0.0909   0.7751   0.0494
  -2.000   0.2247   0.01340   0.00476  -0.0905   0.7654   0.0515
  -1.750   0.2510   0.01314   0.00451  -0.0901   0.7550   0.0549
  -1.500   0.2780   0.01294   0.00427  -0.0897   0.7453   0.0595
  -1.250   0.3045   0.01272   0.00406  -0.0893   0.7347   0.0662
  -1.000   0.3312   0.01252   0.00388  -0.0889   0.7245   0.0804
  -0.750   0.3579   0.01233   0.00375  -0.0886   0.7146   0.1090
  -0.500   0.3841   0.01213   0.00368  -0.0882   0.7038   0.1478
  -0.250   0.4103   0.01193   0.00361  -0.0879   0.6935   0.1981
   0.000   0.4351   0.01151   0.00356  -0.0874   0.6829   0.3191
   0.250   0.4509   0.01035   0.00368  -0.0850   0.6726   0.6923
   0.500   0.5066   0.00996   0.00376  -0.0895   0.6611   0.9793
   0.750   0.5475   0.01003   0.00369  -0.0922   0.6489   1.0000
   1.000   0.5718   0.01013   0.00368  -0.0914   0.6369   1.0000
   1.250   0.5961   0.01024   0.00368  -0.0906   0.6251   1.0000
   1.500   0.6205   0.01037   0.00369  -0.0897   0.6132   1.0000
   1.750   0.6448   0.01050   0.00372  -0.0889   0.6005   1.0000
   2.000   0.6693   0.01064   0.00377  -0.0882   0.5877   1.0000
   2.250   0.6937   0.01079   0.00383  -0.0874   0.5747   1.0000
   2.500   0.7180   0.01095   0.00390  -0.0866   0.5618   1.0000
   2.750   0.7423   0.01113   0.00399  -0.0858   0.5486   1.0000
   3.000   0.7666   0.01131   0.00409  -0.0850   0.5348   1.0000
   3.250   0.7910   0.01150   0.00421  -0.0843   0.5211   1.0000
   3.500   0.8152   0.01170   0.00435  -0.0836   0.5078   1.0000
   3.750   0.8393   0.01192   0.00449  -0.0828   0.4944   1.0000
   4.000   0.8633   0.01216   0.00465  -0.0820   0.4813   1.0000
   4.250   0.8869   0.01242   0.00483  -0.0812   0.4672   1.0000
   4.500   0.9104   0.01268   0.00502  -0.0804   0.4522   1.0000
   4.750   0.9335   0.01296   0.00523  -0.0795   0.4366   1.0000
   5.000   0.9564   0.01326   0.00546  -0.0786   0.4205   1.0000
   5.250   0.9791   0.01357   0.00570  -0.0777   0.4049   1.0000
   5.500   1.0016   0.01389   0.00595  -0.0767   0.3891   1.0000
   5.750   1.0240   0.01422   0.00623  -0.0758   0.3739   1.0000
   6.000   1.0466   0.01455   0.00652  -0.0749   0.3604   1.0000
   6.250   1.0690   0.01489   0.00682  -0.0741   0.3483   1.0000
   6.500   1.0909   0.01525   0.00715  -0.0731   0.3361   1.0000
   6.750   1.1132   0.01559   0.00749  -0.0722   0.3240   1.0000
   7.000   1.1351   0.01595   0.00784  -0.0713   0.3120   1.0000
   7.250   1.1561   0.01634   0.00822  -0.0702   0.2992   1.0000
   7.500   1.1765   0.01676   0.00862  -0.0691   0.2854   1.0000
   7.750   1.1962   0.01721   0.00904  -0.0678   0.2706   1.0000
   8.000   1.2152   0.01769   0.00949  -0.0665   0.2553   1.0000
   8.250   1.2334   0.01820   0.00998  -0.0651   0.2389   1.0000
   8.500   1.2501   0.01878   0.01052  -0.0635   0.2209   1.0000
   8.750   1.2639   0.01944   0.01110  -0.0615   0.2018   1.0000
   9.000   1.2760   0.02021   0.01179  -0.0592   0.1828   1.0000
   9.250   1.2882   0.02103   0.01255  -0.0571   0.1651   1.0000
   9.500   1.2997   0.02191   0.01337  -0.0550   0.1503   1.0000
   9.750   1.3108   0.02284   0.01427  -0.0530   0.1387   1.0000
  10.000   1.3225   0.02376   0.01519  -0.0511   0.1295   1.0000
  10.250   1.3337   0.02473   0.01619  -0.0492   0.1228   1.0000
  10.500   1.3446   0.02574   0.01723  -0.0475   0.1172   1.0000
  10.750   1.3535   0.02692   0.01842  -0.0456   0.1125   1.0000
  11.000   1.3649   0.02796   0.01953  -0.0441   0.1081   1.0000
  11.250   1.3739   0.02921   0.02081  -0.0425   0.1037   1.0000
  11.500   1.3804   0.03067   0.02229  -0.0408   0.1000   1.0000
  11.750   1.3914   0.03184   0.02356  -0.0395   0.0966   1.0000
  12.000   1.4002   0.03322   0.02501  -0.0383   0.0935   1.0000
  12.250   1.4071   0.03479   0.02663  -0.0370   0.0908   1.0000
  12.500   1.4111   0.03664   0.02851  -0.0357   0.0883   1.0000
  12.750   1.4202   0.03811   0.03011  -0.0347   0.0860   1.0000
  13.000   1.4284   0.03971   0.03182  -0.0338   0.0837   1.0000
  13.250   1.4349   0.04147   0.03368  -0.0329   0.0816   1.0000
  13.500   1.4395   0.04346   0.03575  -0.0321   0.0795   1.0000
  13.750   1.4414   0.04576   0.03809  -0.0313   0.0774   1.0000
  14.000   1.4470   0.04775   0.04020  -0.0308   0.0752   1.0000
  14.250   1.4532   0.04973   0.04232  -0.0303   0.0728   1.0000
  14.500   1.4576   0.05195   0.04466  -0.0299   0.0706   1.0000
  14.750   1.4596   0.05450   0.04729  -0.0297   0.0684   1.0000
  15.000   1.4586   0.05746   0.05030  -0.0297   0.0663   1.0000
  15.250   1.4644   0.05970   0.05272  -0.0296   0.0635   1.0000
  15.500   1.4668   0.06240   0.05556  -0.0297   0.0608   1.0000
  15.750   1.4662   0.06556   0.05880  -0.0300   0.0584   1.0000
  16.000   1.4658   0.06876   0.06211  -0.0304   0.0559   1.0000
  16.250   1.4667   0.07182   0.06531  -0.0309   0.0526   1.0000
  16.500   1.4630   0.07561   0.06919  -0.0316   0.0495   1.0000
  16.750   1.4612   0.07918   0.07288  -0.0324   0.0456   1.0000
  17.000   1.4550   0.08347   0.07725  -0.0335   0.0419   1.0000
  17.250   1.4488   0.08779   0.08166  -0.0346   0.0373   1.0000
  17.500   1.4399   0.09259   0.08654  -0.0360   0.0335   1.0000
  17.750   1.4291   0.09780   0.09182  -0.0377   0.0304   1.0000
  18.000   1.4175   0.10322   0.09733  -0.0396   0.0279   1.0000
  18.250   1.4048   0.10890   0.10310  -0.0418   0.0263   1.0000
  18.500   1.3927   0.11458   0.10889  -0.0440   0.0251   1.0000
  18.750   1.3801   0.12047   0.11489  -0.0465   0.0241   1.0000
  19.000   1.3668   0.12661   0.12114  -0.0493   0.0233   1.0000
  19.250   1.3528   0.13302   0.12766  -0.0524   0.0227   1.0000
<< Back to CLARK Z AIRFOIL (clarkz-il)

Polar data table (+)

Polar graphs


<< Back to CLARK Z AIRFOIL (clarkz-il)