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

BOEING 106 AIRFOIL (boe106-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: BOEING 106 AIRFOIL (boe106-il)
Reynolds number: 500,000
Max Cl/Cd: 92.04 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-boe106-il-500000-n5.txt
Download as CSV file: xf-boe106-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 106 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
  ------ -------- --------- --------- -------- -------- --------
 -15.000  -0.9658   0.05131   0.04813  -0.0704   1.0000   0.0142
 -14.750  -0.9970   0.04159   0.03812  -0.0796   1.0000   0.0141
 -14.500  -1.0093   0.03721   0.03353  -0.0823   1.0000   0.0143
 -14.250  -1.0158   0.03434   0.03047  -0.0824   1.0000   0.0145
 -14.000  -1.0187   0.03222   0.02818  -0.0812   1.0000   0.0147
 -13.750  -1.0198   0.03057   0.02638  -0.0790   1.0000   0.0150
 -13.500  -1.0149   0.02945   0.02520  -0.0768   1.0000   0.0153
 -13.250  -1.0057   0.02856   0.02425  -0.0749   1.0000   0.0156
 -13.000  -0.9969   0.02782   0.02345  -0.0726   1.0000   0.0160
 -12.750  -0.9850   0.02706   0.02262  -0.0708   0.9991   0.0164
 -12.500  -0.9586   0.02603   0.02147  -0.0719   0.9949   0.0170
 -12.250  -0.9334   0.02491   0.02019  -0.0727   0.9902   0.0177
 -12.000  -0.9076   0.02377   0.01885  -0.0736   0.9855   0.0183
 -11.750  -0.8813   0.02295   0.01796  -0.0743   0.9801   0.0188
 -11.500  -0.8528   0.02238   0.01734  -0.0751   0.9749   0.0194
 -11.250  -0.8248   0.02184   0.01673  -0.0758   0.9697   0.0200
 -11.000  -0.7979   0.02125   0.01604  -0.0762   0.9631   0.0208
 -10.750  -0.7715   0.02057   0.01523  -0.0765   0.9567   0.0216
 -10.500  -0.7467   0.01992   0.01441  -0.0763   0.9489   0.0222
 -10.250  -0.7227   0.01924   0.01366  -0.0761   0.9415   0.0228
 -10.000  -0.6984   0.01880   0.01317  -0.0757   0.9331   0.0233
  -9.750  -0.6736   0.01842   0.01273  -0.0754   0.9255   0.0240
  -9.500  -0.6494   0.01801   0.01223  -0.0749   0.9170   0.0248
  -9.250  -0.6251   0.01757   0.01168  -0.0743   0.9090   0.0255
  -9.000  -0.6010   0.01710   0.01108  -0.0737   0.8998   0.0262
  -8.750  -0.5765   0.01671   0.01054  -0.0732   0.8901   0.0268
  -8.250  -0.5286   0.01571   0.00944  -0.0720   0.8707   0.0283
  -8.000  -0.5034   0.01536   0.00901  -0.0716   0.8629   0.0290
  -7.750  -0.4780   0.01501   0.00859  -0.0712   0.8546   0.0298
  -7.500  -0.4525   0.01466   0.00814  -0.0708   0.8472   0.0307
  -7.250  -0.4268   0.01430   0.00769  -0.0704   0.8392   0.0314
  -7.000  -0.4010   0.01399   0.00727  -0.0700   0.8319   0.0319
  -6.750  -0.3760   0.01349   0.00674  -0.0696   0.8242   0.0328
  -6.500  -0.3506   0.01313   0.00632  -0.0691   0.8168   0.0337
  -6.250  -0.3245   0.01281   0.00596  -0.0688   0.8095   0.0344
  -6.000  -0.2984   0.01252   0.00562  -0.0685   0.8021   0.0352
  -5.750  -0.2721   0.01225   0.00529  -0.0682   0.7949   0.0361
  -5.500  -0.2457   0.01201   0.00499  -0.0679   0.7873   0.0371
  -5.250  -0.2191   0.01178   0.00470  -0.0676   0.7805   0.0380
  -5.000  -0.1928   0.01151   0.00437  -0.0672   0.7728   0.0388
  -4.750  -0.1668   0.01123   0.00405  -0.0669   0.7657   0.0401
  -4.500  -0.1401   0.01100   0.00380  -0.0666   0.7580   0.0415
  -4.250  -0.1134   0.01083   0.00357  -0.0663   0.7508   0.0430
  -4.000  -0.0864   0.01066   0.00336  -0.0661   0.7428   0.0448
  -3.750  -0.0597   0.01051   0.00315  -0.0658   0.7350   0.0469
  -3.500  -0.0329   0.01031   0.00295  -0.0656   0.7266   0.0502
  -3.250  -0.0061   0.01018   0.00277  -0.0653   0.7182   0.0537
  -3.000   0.0206   0.01002   0.00261  -0.0650   0.7077   0.0593
  -2.750   0.0470   0.00987   0.00246  -0.0647   0.6961   0.0700
  -2.500   0.0730   0.00970   0.00234  -0.0643   0.6843   0.0935
  -2.250   0.0996   0.00957   0.00225  -0.0641   0.6732   0.1147
  -2.000   0.1264   0.00947   0.00216  -0.0638   0.6632   0.1317
  -1.750   0.1531   0.00938   0.00207  -0.0636   0.6533   0.1476
  -1.500   0.1797   0.00928   0.00200  -0.0633   0.6423   0.1662
  -1.250   0.2054   0.00907   0.00192  -0.0630   0.6316   0.2153
  -1.000   0.2301   0.00878   0.00187  -0.0625   0.6212   0.3006
  -0.750   0.2552   0.00854   0.00184  -0.0621   0.6105   0.3737
  -0.500   0.2792   0.00822   0.00180  -0.0614   0.5999   0.4710
  -0.250   0.3009   0.00783   0.00181  -0.0603   0.5892   0.6007
   0.000   0.3239   0.00762   0.00185  -0.0592   0.5782   0.6947
   0.250   0.3486   0.00753   0.00189  -0.0584   0.5685   0.7460
   0.500   0.3732   0.00747   0.00193  -0.0576   0.5592   0.7921
   1.000   0.4255   0.00747   0.00207  -0.0564   0.5407   0.8679
   1.250   0.4548   0.00754   0.00216  -0.0565   0.5309   0.9034
   1.500   0.4873   0.00765   0.00226  -0.0574   0.5216   0.9296
   1.750   0.5238   0.00780   0.00237  -0.0591   0.5124   0.9520
   2.000   0.5604   0.00793   0.00247  -0.0610   0.5037   0.9693
   2.250   0.5998   0.00809   0.00257  -0.0634   0.4949   0.9801
   2.500   0.6416   0.00821   0.00266  -0.0665   0.4856   0.9877
   2.750   0.6776   0.00835   0.00274  -0.0684   0.4759   0.9926
   3.000   0.7112   0.00850   0.00283  -0.0698   0.4625   0.9955
   3.500   0.7747   0.00880   0.00300  -0.0719   0.4353   1.0000
   3.750   0.7974   0.00894   0.00311  -0.0709   0.4243   1.0000
   4.000   0.8196   0.00911   0.00323  -0.0699   0.4128   1.0000
   4.500   0.8640   0.00946   0.00349  -0.0678   0.3897   1.0000
   4.750   0.8861   0.00964   0.00364  -0.0667   0.3796   1.0000
   5.000   0.9075   0.00986   0.00380  -0.0656   0.3679   1.0000
   5.250   0.9283   0.01011   0.00399  -0.0643   0.3525   1.0000
   5.500   0.9490   0.01038   0.00418  -0.0630   0.3378   1.0000
   5.750   0.9703   0.01061   0.00438  -0.0619   0.3251   1.0000
   6.000   0.9916   0.01085   0.00459  -0.0607   0.3130   1.0000
   6.250   1.0129   0.01111   0.00481  -0.0596   0.3013   1.0000
   6.500   1.0341   0.01139   0.00505  -0.0584   0.2891   1.0000
   6.750   1.0553   0.01168   0.00530  -0.0573   0.2757   1.0000
   7.000   1.0762   0.01200   0.00557  -0.0562   0.2584   1.0000
   7.250   1.0946   0.01248   0.00592  -0.0548   0.2299   1.0000
   7.500   1.1089   0.01322   0.00642  -0.0527   0.1880   1.0000
   7.750   1.1251   0.01386   0.00693  -0.0510   0.1643   1.0000
   8.000   1.1434   0.01436   0.00738  -0.0496   0.1512   1.0000
   8.250   1.1621   0.01483   0.00782  -0.0483   0.1410   1.0000
   8.500   1.1798   0.01532   0.00828  -0.0468   0.1316   1.0000
   8.750   1.1977   0.01573   0.00871  -0.0453   0.1237   1.0000
   9.000   1.2138   0.01625   0.00920  -0.0437   0.1155   1.0000
   9.250   1.2306   0.01674   0.00969  -0.0421   0.1078   1.0000
   9.500   1.2465   0.01730   0.01024  -0.0406   0.1006   1.0000
   9.750   1.2617   0.01792   0.01085  -0.0390   0.0927   1.0000
  10.000   1.2771   0.01855   0.01147  -0.0375   0.0857   1.0000
  10.500   1.3062   0.01995   0.01288  -0.0346   0.0747   1.0000
  10.750   1.3197   0.02074   0.01368  -0.0331   0.0701   1.0000
  11.000   1.3331   0.02157   0.01453  -0.0318   0.0661   1.0000
  11.250   1.3467   0.02240   0.01540  -0.0305   0.0625   1.0000
  11.500   1.3587   0.02338   0.01639  -0.0291   0.0594   1.0000
  11.750   1.3707   0.02438   0.01744  -0.0279   0.0569   1.0000
  12.000   1.3836   0.02536   0.01847  -0.0268   0.0543   1.0000
  12.250   1.3945   0.02650   0.01965  -0.0257   0.0514   1.0000
  12.500   1.4035   0.02783   0.02100  -0.0245   0.0487   1.0000
  12.750   1.4147   0.02902   0.02226  -0.0236   0.0469   1.0000
  13.000   1.4248   0.03032   0.02362  -0.0226   0.0446   1.0000
  13.250   1.4329   0.03183   0.02518  -0.0217   0.0420   1.0000
  13.500   1.4405   0.03343   0.02683  -0.0208   0.0396   1.0000
  13.750   1.4477   0.03508   0.02854  -0.0201   0.0365   1.0000
  14.000   1.4526   0.03700   0.03049  -0.0193   0.0333   1.0000
  14.250   1.4564   0.03906   0.03259  -0.0186   0.0291   1.0000
  14.500   1.4578   0.04141   0.03496  -0.0179   0.0247   1.0000
  14.750   1.4570   0.04407   0.03764  -0.0174   0.0207   1.0000
  15.000   1.4557   0.04688   0.04050  -0.0170   0.0180   1.0000
  15.250   1.4537   0.04984   0.04352  -0.0168   0.0162   1.0000
  15.500   1.4526   0.05280   0.04657  -0.0167   0.0150   1.0000
  15.750   1.4509   0.05592   0.04978  -0.0168   0.0141   1.0000
  16.000   1.4479   0.05927   0.05321  -0.0169   0.0133   1.0000
  16.250   1.4439   0.06280   0.05682  -0.0173   0.0127   1.0000
  16.500   1.4419   0.06617   0.06029  -0.0177   0.0122   1.0000
  16.750   1.4387   0.06975   0.06398  -0.0182   0.0118   1.0000
  17.000   1.4342   0.07354   0.06788  -0.0189   0.0114   1.0000
  17.250   1.4286   0.07754   0.07197  -0.0197   0.0110   1.0000
  17.500   1.4220   0.08175   0.07628  -0.0207   0.0107   1.0000
  17.750   1.4143   0.08620   0.08084  -0.0218   0.0104   1.0000
  18.000   1.4052   0.09093   0.08567  -0.0231   0.0102   1.0000
<< Back to BOEING 106 AIRFOIL (boe106-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 106 AIRFOIL (boe106-il)