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

CAST 10-2/DOA 2 AIRFOIL (cast102-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: CAST 10-2/DOA 2 AIRFOIL (cast102-il)
Reynolds number: 500,000
Max Cl/Cd: 66.8 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-cast102-il-500000-n5.txt
Download as CSV file: xf-cast102-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CAST 10-2/DOA 2 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5048   0.07849   0.07610  -0.0616   0.9896   0.0099
 -10.000  -0.5229   0.06909   0.06660  -0.0709   0.9820   0.0097
  -9.750  -0.5402   0.06250   0.05985  -0.0759   0.9730   0.0097
  -9.500  -0.5560   0.05785   0.05504  -0.0763   0.9623   0.0096
  -9.250  -0.5626   0.05344   0.05043  -0.0760   0.9551   0.0097
  -9.000  -0.5650   0.04928   0.04605  -0.0751   0.9483   0.0097
  -8.750  -0.5629   0.04522   0.04173  -0.0740   0.9436   0.0098
  -8.500  -0.5579   0.04126   0.03750  -0.0727   0.9384   0.0099
  -8.250  -0.5500   0.03735   0.03329  -0.0712   0.9338   0.0100
  -8.000  -0.5385   0.03376   0.02936  -0.0697   0.9301   0.0103
  -7.750  -0.5256   0.02959   0.02477  -0.0681   0.9265   0.0108
  -7.500  -0.5105   0.02553   0.02024  -0.0665   0.9229   0.0110
  -7.250  -0.4916   0.02204   0.01623  -0.0651   0.9199   0.0114
  -7.000  -0.4693   0.02014   0.01404  -0.0645   0.9173   0.0117
  -6.750  -0.4454   0.01909   0.01287  -0.0642   0.9149   0.0119
  -6.500  -0.4205   0.01827   0.01195  -0.0640   0.9126   0.0122
  -6.250  -0.3953   0.01749   0.01109  -0.0639   0.9100   0.0125
  -6.000  -0.3698   0.01672   0.01023  -0.0637   0.9075   0.0128
  -5.750  -0.3443   0.01602   0.00944  -0.0635   0.9051   0.0132
  -5.500  -0.3186   0.01538   0.00872  -0.0633   0.9027   0.0139
  -5.250  -0.2930   0.01474   0.00799  -0.0630   0.9005   0.0144
  -5.000  -0.2673   0.01415   0.00730  -0.0628   0.8986   0.0148
  -4.750  -0.2427   0.01341   0.00654  -0.0625   0.8963   0.0152
  -4.500  -0.2173   0.01291   0.00603  -0.0624   0.8939   0.0156
  -4.250  -0.1912   0.01250   0.00562  -0.0623   0.8914   0.0162
  -4.000  -0.1649   0.01212   0.00521  -0.0623   0.8890   0.0167
  -3.750  -0.1384   0.01175   0.00483  -0.0623   0.8868   0.0174
  -3.500  -0.1112   0.01148   0.00452  -0.0623   0.8850   0.0184
  -3.250  -0.0842   0.01113   0.00414  -0.0624   0.8833   0.0193
  -3.000  -0.0570   0.01083   0.00385  -0.0626   0.8812   0.0202
  -2.750  -0.0295   0.01060   0.00362  -0.0627   0.8787   0.0213
  -2.500  -0.0017   0.01039   0.00340  -0.0629   0.8763   0.0226
  -2.250   0.0264   0.01021   0.00321  -0.0631   0.8741   0.0240
  -2.000   0.0544   0.01001   0.00301  -0.0634   0.8722   0.0272
  -1.750   0.0828   0.00986   0.00284  -0.0636   0.8704   0.0312
  -1.250   0.1315   0.00747   0.00234  -0.0643   0.8662   0.5975
  -1.000   0.1593   0.00741   0.00238  -0.0643   0.8626   0.6374
  -0.750   0.1873   0.00735   0.00235  -0.0643   0.8572   0.6577
  -0.500   0.2154   0.00729   0.00229  -0.0642   0.8507   0.6731
  -0.250   0.2432   0.00725   0.00226  -0.0642   0.8429   0.6842
   0.000   0.2714   0.00721   0.00220  -0.0642   0.8355   0.6932
   0.250   0.2995   0.00717   0.00212  -0.0642   0.8252   0.6975
   0.500   0.3274   0.00713   0.00207  -0.0642   0.8138   0.7005
   0.750   0.3553   0.00710   0.00202  -0.0642   0.8011   0.7038
   1.000   0.3833   0.00709   0.00198  -0.0642   0.7876   0.7074
   1.250   0.4115   0.00710   0.00197  -0.0643   0.7750   0.7113
   1.500   0.4392   0.00712   0.00196  -0.0642   0.7565   0.7148
   1.750   0.4661   0.00717   0.00194  -0.0640   0.7288   0.7185
   2.000   0.4910   0.00735   0.00194  -0.0633   0.6794   0.7225
   2.250   0.5083   0.00804   0.00211  -0.0613   0.5682   0.7265
   2.500   0.5239   0.00893   0.00246  -0.0591   0.4469   0.7304
   2.750   0.5439   0.00955   0.00274  -0.0579   0.3664   0.7348
   3.000   0.5673   0.00995   0.00294  -0.0573   0.3212   0.7397
   3.500   0.6167   0.01052   0.00332  -0.0565   0.2620   0.7487
   3.750   0.6414   0.01081   0.00352  -0.0561   0.2354   0.7536
   4.000   0.6660   0.01112   0.00373  -0.0557   0.2108   0.7588
   4.250   0.6905   0.01139   0.00394  -0.0553   0.1897   0.7640
   4.500   0.7152   0.01166   0.00416  -0.0549   0.1722   0.7700
   4.750   0.7400   0.01193   0.00439  -0.0545   0.1574   0.7759
   5.000   0.7646   0.01218   0.00463  -0.0541   0.1447   0.7818
   5.250   0.7894   0.01243   0.00488  -0.0538   0.1337   0.7884
   5.500   0.8139   0.01268   0.00513  -0.0533   0.1247   0.7950
   5.750   0.8382   0.01295   0.00540  -0.0529   0.1144   0.8029
   6.000   0.8620   0.01324   0.00568  -0.0523   0.1016   0.8109
   6.250   0.8857   0.01354   0.00597  -0.0518   0.0904   0.8199
   6.500   0.9088   0.01385   0.00628  -0.0511   0.0796   0.8292
   6.750   0.9315   0.01420   0.00663  -0.0504   0.0679   0.8397
   7.750   1.0108   0.01594   0.00843  -0.0450   0.0259   0.9162
   8.000   1.0310   0.01632   0.00889  -0.0437   0.0236   0.9757
   8.250   1.0535   0.01678   0.00938  -0.0431   0.0221   1.0000
   8.500   1.0754   0.01725   0.00991  -0.0424   0.0211   1.0000
   8.750   1.0966   0.01776   0.01047  -0.0415   0.0201   1.0000
   9.000   1.1162   0.01832   0.01107  -0.0403   0.0192   1.0000
   9.250   1.1342   0.01900   0.01179  -0.0390   0.0183   1.0000
   9.500   1.1519   0.01968   0.01254  -0.0375   0.0177   1.0000
   9.750   1.1708   0.02024   0.01317  -0.0363   0.0171   1.0000
  10.000   1.1890   0.02085   0.01385  -0.0350   0.0165   1.0000
  10.250   1.2065   0.02151   0.01458  -0.0337   0.0159   1.0000
  10.500   1.2233   0.02222   0.01535  -0.0323   0.0154   1.0000
  10.750   1.2392   0.02298   0.01619  -0.0308   0.0150   1.0000
  11.000   1.2539   0.02383   0.01710  -0.0292   0.0146   1.0000
  11.250   1.2661   0.02486   0.01819  -0.0273   0.0142   1.0000
  11.500   1.2759   0.02605   0.01947  -0.0251   0.0138   1.0000
  11.750   1.2893   0.02698   0.02050  -0.0235   0.0136   1.0000
  12.000   1.3016   0.02799   0.02163  -0.0218   0.0134   1.0000
  12.250   1.3125   0.02910   0.02284  -0.0200   0.0131   1.0000
  12.500   1.3227   0.03028   0.02413  -0.0183   0.0129   1.0000
  12.750   1.3321   0.03152   0.02548  -0.0165   0.0126   1.0000
  13.000   1.3410   0.03280   0.02687  -0.0148   0.0123   1.0000
  13.250   1.3492   0.03416   0.02833  -0.0131   0.0120   1.0000
  13.500   1.3563   0.03561   0.02989  -0.0115   0.0118   1.0000
  13.750   1.3623   0.03719   0.03157  -0.0099   0.0116   1.0000
  14.000   1.3667   0.03895   0.03345  -0.0084   0.0114   1.0000
  14.250   1.3695   0.04089   0.03549  -0.0069   0.0112   1.0000
  14.500   1.3702   0.04308   0.03780  -0.0055   0.0111   1.0000
  14.750   1.3687   0.04559   0.04042  -0.0043   0.0110   1.0000
  15.000   1.3642   0.04850   0.04346  -0.0032   0.0108   1.0000
  15.250   1.3565   0.05191   0.04701  -0.0024   0.0107   1.0000
  15.500   1.3502   0.05543   0.05067  -0.0021   0.0106   1.0000
  15.750   1.3441   0.05919   0.05460  -0.0024   0.0105   1.0000
  16.000   1.3364   0.06343   0.05901  -0.0032   0.0105   1.0000
  16.250   1.3264   0.06837   0.06412  -0.0047   0.0104   1.0000
  16.500   1.3149   0.07397   0.06989  -0.0070   0.0104   1.0000
  16.750   1.3010   0.08033   0.07643  -0.0100   0.0103   1.0000
  17.000   1.2852   0.08739   0.08367  -0.0137   0.0103   1.0000
  17.250   1.2675   0.09507   0.09151  -0.0178   0.0103   1.0000
  17.500   1.2482   0.10317   0.09978  -0.0222   0.0103   1.0000
  17.750   1.2278   0.11150   0.10826  -0.0267   0.0102   1.0000
  18.000   1.2068   0.12009   0.11699  -0.0314   0.0102   1.0000
<< Back to CAST 10-2/DOA 2 AIRFOIL (cast102-il)

Polar data table (+)

Polar graphs


<< Back to CAST 10-2/DOA 2 AIRFOIL (cast102-il)