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

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


Details Polar file
Airfoil: CLARK Y AIRFOIL (clarky-il)
Reynolds number: 200,000
Max Cl/Cd: 69.66 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-clarky-il-200000-n5.txt
Download as CSV file: xf-clarky-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CLARK Y 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3991   0.10855   0.10470  -0.0373   1.0000   0.0283
 -10.750  -0.4131   0.10067   0.09685  -0.0415   1.0000   0.0292
 -10.500  -0.4187   0.09546   0.09168  -0.0438   1.0000   0.0297
 -10.250  -0.4148   0.09285   0.08910  -0.0444   1.0000   0.0300
 -10.000  -0.4155   0.08932   0.08560  -0.0456   1.0000   0.0303
  -9.750  -0.4199   0.08535   0.08167  -0.0469   1.0000   0.0305
  -9.500  -0.4281   0.08133   0.07771  -0.0480   1.0000   0.0308
  -9.250  -0.4414   0.07744   0.07389  -0.0485   1.0000   0.0311
  -9.000  -0.5389   0.05070   0.04695  -0.0702   0.9886   0.0307
  -8.750  -0.5278   0.04461   0.04051  -0.0762   0.9817   0.0316
  -8.500  -0.5181   0.03875   0.03412  -0.0798   0.9735   0.0325
  -8.250  -0.5015   0.03390   0.02868  -0.0823   0.9677   0.0333
  -8.000  -0.4817   0.03065   0.02489  -0.0834   0.9610   0.0344
  -7.750  -0.4552   0.02808   0.02177  -0.0849   0.9570   0.0354
  -7.500  -0.4319   0.02623   0.01955  -0.0850   0.9504   0.0358
  -7.250  -0.4054   0.02437   0.01748  -0.0857   0.9452   0.0364
  -7.000  -0.3749   0.02307   0.01602  -0.0868   0.9416   0.0371
  -6.750  -0.3498   0.02218   0.01503  -0.0867   0.9344   0.0380
  -6.500  -0.3207   0.02126   0.01397  -0.0872   0.9291   0.0390
  -6.250  -0.2910   0.02029   0.01282  -0.0878   0.9244   0.0400
  -6.000  -0.2654   0.01941   0.01179  -0.0875   0.9166   0.0407
  -5.750  -0.2356   0.01854   0.01078  -0.0879   0.9115   0.0414
  -5.500  -0.2093   0.01783   0.00994  -0.0876   0.9041   0.0422
  -5.250  -0.1811   0.01716   0.00916  -0.0877   0.8977   0.0429
  -5.000  -0.1541   0.01645   0.00841  -0.0876   0.8910   0.0440
  -4.750  -0.1275   0.01590   0.00785  -0.0874   0.8834   0.0455
  -4.500  -0.0995   0.01543   0.00733  -0.0873   0.8772   0.0470
  -4.250  -0.0733   0.01500   0.00686  -0.0870   0.8689   0.0485
  -4.000  -0.0452   0.01457   0.00636  -0.0869   0.8626   0.0501
  -3.750  -0.0191   0.01423   0.00595  -0.0865   0.8540   0.0517
  -3.500   0.0087   0.01382   0.00552  -0.0864   0.8476   0.0543
  -3.250   0.0349   0.01355   0.00523  -0.0860   0.8389   0.0583
  -3.000   0.0628   0.01325   0.00490  -0.0859   0.8319   0.0644
  -2.750   0.0892   0.01299   0.00464  -0.0855   0.8222   0.0746
  -2.500   0.1161   0.01269   0.00435  -0.0851   0.8128   0.0905
  -2.250   0.1430   0.01240   0.00410  -0.0848   0.8027   0.1120
  -2.000   0.1691   0.01213   0.00393  -0.0845   0.7918   0.1432
  -1.750   0.1955   0.01185   0.00377  -0.0841   0.7816   0.1908
  -1.500   0.2217   0.01157   0.00365  -0.0838   0.7712   0.2491
  -1.250   0.2477   0.01135   0.00356  -0.0834   0.7593   0.2987
  -1.000   0.2732   0.01108   0.00346  -0.0829   0.7476   0.3611
  -0.750   0.2976   0.01069   0.00339  -0.0822   0.7373   0.4637
  -0.500   0.3200   0.01022   0.00340  -0.0810   0.7273   0.6001
  -0.250   0.3414   0.00982   0.00345  -0.0791   0.7176   0.7430
   0.250   0.4236   0.00964   0.00350  -0.0832   0.6966   0.9672
   0.500   0.4701   0.00968   0.00344  -0.0870   0.6857   0.9947
   0.750   0.5017   0.00975   0.00338  -0.0878   0.6743   1.0000
   1.000   0.5260   0.00983   0.00337  -0.0871   0.6622   1.0000
   1.250   0.5503   0.00992   0.00337  -0.0863   0.6495   1.0000
   1.500   0.5745   0.01003   0.00339  -0.0855   0.6371   1.0000
   1.750   0.5988   0.01014   0.00342  -0.0847   0.6246   1.0000
   2.000   0.6230   0.01026   0.00345  -0.0839   0.6112   1.0000
   2.250   0.6471   0.01040   0.00350  -0.0831   0.5965   1.0000
   2.500   0.6711   0.01055   0.00357  -0.0822   0.5809   1.0000
   2.750   0.6951   0.01071   0.00365  -0.0814   0.5649   1.0000
   3.000   0.7190   0.01089   0.00374  -0.0806   0.5489   1.0000
   3.250   0.7429   0.01109   0.00386  -0.0797   0.5320   1.0000
   3.500   0.7664   0.01130   0.00399  -0.0789   0.5138   1.0000
   3.750   0.7898   0.01154   0.00414  -0.0780   0.4947   1.0000
   4.000   0.8129   0.01180   0.00431  -0.0770   0.4759   1.0000
   4.250   0.8361   0.01207   0.00450  -0.0761   0.4573   1.0000
   4.500   0.8591   0.01235   0.00471  -0.0753   0.4385   1.0000
   4.750   0.8819   0.01266   0.00494  -0.0743   0.4206   1.0000
   5.000   0.9045   0.01299   0.00519  -0.0734   0.4038   1.0000
   5.250   0.9268   0.01334   0.00546  -0.0724   0.3869   1.0000
   5.500   0.9488   0.01372   0.00575  -0.0715   0.3708   1.0000
   5.750   0.9706   0.01410   0.00606  -0.0705   0.3551   1.0000
   6.000   0.9928   0.01448   0.00638  -0.0696   0.3413   1.0000
   6.250   1.0154   0.01484   0.00671  -0.0687   0.3307   1.0000
   6.500   1.0378   0.01521   0.00707  -0.0679   0.3218   1.0000
   6.750   1.0610   0.01553   0.00742  -0.0671   0.3128   1.0000
   7.000   1.0829   0.01592   0.00780  -0.0662   0.3043   1.0000
   7.250   1.1060   0.01623   0.00817  -0.0655   0.2958   1.0000
   7.500   1.1279   0.01661   0.00858  -0.0646   0.2874   1.0000
   7.750   1.1498   0.01697   0.00898  -0.0637   0.2774   1.0000
   8.000   1.1714   0.01733   0.00939  -0.0627   0.2667   1.0000
   8.250   1.1917   0.01774   0.00982  -0.0616   0.2529   1.0000
   8.500   1.2108   0.01819   0.01027  -0.0603   0.2355   1.0000
   8.750   1.2278   0.01875   0.01076  -0.0588   0.2136   1.0000
   9.000   1.2423   0.01941   0.01133  -0.0569   0.1887   1.0000
   9.250   1.2544   0.02020   0.01201  -0.0547   0.1662   1.0000
   9.500   1.2651   0.02111   0.01283  -0.0524   0.1462   1.0000
   9.750   1.2765   0.02203   0.01368  -0.0503   0.1304   1.0000
  10.000   1.2881   0.02296   0.01458  -0.0483   0.1180   1.0000
  10.250   1.3006   0.02384   0.01547  -0.0465   0.1083   1.0000
  10.500   1.3127   0.02476   0.01642  -0.0447   0.1007   1.0000
  10.750   1.3231   0.02581   0.01749  -0.0429   0.0931   1.0000
  11.000   1.3348   0.02680   0.01854  -0.0413   0.0860   1.0000
  11.250   1.3442   0.02799   0.01975  -0.0396   0.0794   1.0000
  11.500   1.3548   0.02911   0.02093  -0.0381   0.0716   1.0000
  11.750   1.3635   0.03042   0.02226  -0.0365   0.0620   1.0000
  12.000   1.3695   0.03199   0.02383  -0.0349   0.0497   1.0000
  12.250   1.3701   0.03406   0.02581  -0.0331   0.0327   1.0000
  12.500   1.3679   0.03648   0.02816  -0.0314   0.0246   1.0000
  12.750   1.3678   0.03881   0.03055  -0.0300   0.0216   1.0000
  13.000   1.3679   0.04122   0.03304  -0.0289   0.0199   1.0000
  13.250   1.3663   0.04388   0.03579  -0.0279   0.0187   1.0000
  13.500   1.3671   0.04640   0.03845  -0.0271   0.0179   1.0000
  13.750   1.3662   0.04917   0.04136  -0.0266   0.0172   1.0000
  14.000   1.3640   0.05219   0.04453  -0.0262   0.0167   1.0000
  14.250   1.3602   0.05552   0.04801  -0.0261   0.0162   1.0000
  14.500   1.3546   0.05920   0.05183  -0.0262   0.0159   1.0000
  14.750   1.3474   0.06326   0.05603  -0.0267   0.0156   1.0000
  15.000   1.3382   0.06776   0.06068  -0.0275   0.0153   1.0000
  15.250   1.3286   0.07247   0.06552  -0.0285   0.0151   1.0000
  15.500   1.3212   0.07701   0.07022  -0.0296   0.0149   1.0000
  15.750   1.3128   0.08183   0.07520  -0.0310   0.0147   1.0000
  16.000   1.3038   0.08685   0.08037  -0.0325   0.0145   1.0000
  16.250   1.2942   0.09206   0.08572  -0.0342   0.0143   1.0000
  16.500   1.2845   0.09738   0.09118  -0.0361   0.0141   1.0000
<< Back to CLARK Y AIRFOIL (clarky-il)

Polar data table (+)

Polar graphs


<< Back to CLARK Y AIRFOIL (clarky-il)