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

CLARK W AIRFOIL (clarkw-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: CLARK W AIRFOIL (clarkw-il)
Reynolds number: 50,000
Max Cl/Cd: 32.22 at α=9.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-clarkw-il-50000.txt
Download as CSV file: xf-clarkw-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CLARK W AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3420   0.10855   0.10124  -0.0271   1.0000   0.2455
  -8.500  -0.3453   0.10676   0.09954  -0.0262   1.0000   0.2544
  -8.250  -0.3687   0.10618   0.09913  -0.0254   1.0000   0.2582
  -8.000  -0.3437   0.10153   0.09444  -0.0237   1.0000   0.2685
  -7.750  -0.3824   0.10263   0.09578  -0.0215   1.0000   0.2734
  -7.500  -0.3524   0.09709   0.09020  -0.0203   1.0000   0.2813
  -7.250  -0.3781   0.09688   0.09014  -0.0173   1.0000   0.2887
  -7.000  -0.3967   0.09527   0.08869  -0.0143   1.0000   0.2925
  -6.750  -0.3832   0.09191   0.08534  -0.0123   1.0000   0.3010
  -6.500  -0.4349   0.09319   0.08684  -0.0097   1.0000   0.3072
  -6.250  -0.4098   0.08835   0.08200  -0.0067   1.0000   0.3135
  -6.000  -0.4443   0.08831   0.08211  -0.0054   1.0000   0.3237
  -5.500  -0.4541   0.08326   0.07719  -0.0018   1.0000   0.3414
  -5.250  -0.4409   0.07967   0.07363   0.0018   1.0000   0.3482
  -4.500  -0.4163   0.05481   0.04731  -0.0317   1.0000   0.1768
  -4.250  -0.3999   0.05166   0.04405  -0.0315   1.0000   0.1737
  -4.000  -0.3801   0.04782   0.03993  -0.0326   1.0000   0.1690
  -3.750  -0.3495   0.04165   0.03253  -0.0366   1.0000   0.1604
  -3.500  -0.3255   0.03888   0.02936  -0.0371   1.0000   0.1610
  -3.250  -0.3048   0.03747   0.02800  -0.0366   1.0000   0.1656
  -3.000  -0.2806   0.03576   0.02592  -0.0368   1.0000   0.1723
  -2.750  -0.2553   0.03399   0.02378  -0.0370   1.0000   0.1793
  -2.500  -0.2309   0.03273   0.02229  -0.0369   1.0000   0.1894
  -2.250  -0.2069   0.03178   0.02117  -0.0367   1.0000   0.2068
  -2.000  -0.1843   0.03111   0.02061  -0.0363   1.0000   0.2314
  -1.750  -0.1622   0.03073   0.02028  -0.0357   1.0000   0.2672
  -1.500  -0.1424   0.03072   0.02035  -0.0350   1.0000   0.3103
  -1.250  -0.1227   0.03066   0.02040  -0.0343   1.0000   0.3540
  -1.000  -0.1003   0.03026   0.02002  -0.0339   1.0000   0.3936
  -0.750  -0.0753   0.02988   0.01959  -0.0341   1.0000   0.4393
  -0.500  -0.0496   0.02938   0.01930  -0.0344   1.0000   0.4942
  -0.250  -0.0228   0.02872   0.01917  -0.0347   1.0000   0.5757
   0.000  -0.0080   0.02750   0.01927  -0.0326   0.9989   1.0000
   0.250   0.0583   0.02916   0.02014  -0.0413   0.9833   1.0000
   0.500   0.1083   0.03037   0.02096  -0.0467   0.9663   1.0000
   0.750   0.1546   0.03146   0.02176  -0.0513   0.9490   1.0000
   1.000   0.1997   0.03246   0.02254  -0.0554   0.9316   1.0000
   1.250   0.2437   0.03338   0.02328  -0.0591   0.9142   1.0000
   1.500   0.2870   0.03420   0.02396  -0.0625   0.8969   1.0000
   1.750   0.3297   0.03493   0.02458  -0.0656   0.8796   1.0000
   2.000   0.3715   0.03555   0.02512  -0.0684   0.8626   1.0000
   2.250   0.4064   0.03614   0.02565  -0.0699   0.8446   1.0000
   2.500   0.4373   0.03673   0.02621  -0.0707   0.8262   1.0000
   2.750   0.4698   0.03726   0.02672  -0.0716   0.8084   1.0000
   3.000   0.5028   0.03775   0.02721  -0.0724   0.7912   1.0000
   3.250   0.5358   0.03818   0.02764  -0.0731   0.7744   1.0000
   3.500   0.5682   0.03860   0.02807  -0.0736   0.7582   1.0000
   3.750   0.6002   0.03895   0.02846  -0.0740   0.7423   1.0000
   4.000   0.6323   0.03927   0.02880  -0.0742   0.7271   1.0000
   4.250   0.6695   0.03928   0.02887  -0.0748   0.7134   1.0000
   4.500   0.7150   0.03874   0.02842  -0.0761   0.7014   1.0000
   4.750   0.7387   0.03931   0.02903  -0.0752   0.6863   1.0000
   5.000   0.7615   0.03993   0.02972  -0.0741   0.6714   1.0000
   5.250   0.7824   0.04073   0.03060  -0.0730   0.6569   1.0000
   5.500   0.8036   0.04154   0.03148  -0.0720   0.6429   1.0000
   5.750   0.8286   0.04213   0.03214  -0.0712   0.6299   1.0000
   6.000   0.8836   0.04058   0.03075  -0.0723   0.6207   1.0000
   6.250   0.8958   0.04200   0.03225  -0.0705   0.6061   1.0000
   6.500   0.9059   0.04368   0.03402  -0.0686   0.5919   1.0000
   6.750   0.9201   0.04504   0.03549  -0.0670   0.5780   1.0000
   7.000   0.9449   0.04549   0.03606  -0.0658   0.5646   1.0000
   7.250   0.9847   0.04444   0.03514  -0.0650   0.5497   1.0000
   7.500   1.0235   0.04334   0.03419  -0.0641   0.5333   1.0000
   7.750   1.0582   0.04266   0.03364  -0.0630   0.5165   1.0000
   8.000   1.0910   0.04220   0.03329  -0.0619   0.4990   1.0000
   8.250   1.1326   0.04104   0.03224  -0.0613   0.4792   1.0000
   8.500   1.1694   0.04008   0.03129  -0.0602   0.4562   1.0000
   8.750   1.2103   0.03849   0.02958  -0.0592   0.4266   1.0000
   9.000   1.2199   0.03894   0.03012  -0.0557   0.3997   1.0000
   9.250   1.2380   0.03877   0.02988  -0.0528   0.3688   1.0000
   9.500   1.2519   0.03886   0.02985  -0.0495   0.3363   1.0000
   9.750   1.2614   0.03925   0.03008  -0.0459   0.3030   1.0000
  10.000   1.2663   0.04007   0.03074  -0.0419   0.2713   1.0000
  10.250   1.2683   0.04137   0.03191  -0.0380   0.2432   1.0000
  10.500   1.2745   0.04282   0.03313  -0.0347   0.2188   1.0000
  10.750   1.2786   0.04474   0.03504  -0.0316   0.1996   1.0000
  11.000   1.2867   0.04673   0.03699  -0.0291   0.1836   1.0000
  11.250   1.2954   0.04900   0.03930  -0.0268   0.1710   1.0000
  11.500   1.3089   0.05147   0.04177  -0.0253   0.1601   1.0000
  11.750   1.3312   0.05401   0.04417  -0.0249   0.1493   1.0000
  12.000   1.3182   0.05709   0.04772  -0.0208   0.1459   1.0000
  12.250   1.3142   0.06020   0.05109  -0.0181   0.1413   1.0000
  12.500   1.3298   0.06329   0.05415  -0.0173   0.1348   1.0000
  12.750   1.3075   0.06704   0.05827  -0.0138   0.1337   1.0000
  13.000   1.2825   0.07134   0.06291  -0.0112   0.1330   1.0000
  13.250   1.2537   0.07637   0.06824  -0.0096   0.1329   1.0000
  13.500   1.2205   0.08236   0.07449  -0.0092   0.1334   1.0000
  13.750   1.1847   0.08948   0.08183  -0.0103   0.1344   1.0000
  14.000   1.1490   0.09775   0.09024  -0.0128   0.1354   1.0000
<< Back to CLARK W AIRFOIL (clarkw-il)

Polar data table (+)

Polar graphs


<< Back to CLARK W AIRFOIL (clarkw-il)