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

BOEING 707 .99 SPAN AIRFOIL (b707e-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: BOEING 707 .99 SPAN AIRFOIL (b707e-il)
Reynolds number: 500,000
Max Cl/Cd: 65.52 at α=1.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-b707e-il-500000.txt
Download as CSV file: xf-b707e-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 707 .99 SPAN 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4542   0.08631   0.08409  -0.0335   1.0000   0.0162
  -8.500  -0.4576   0.08209   0.07989  -0.0368   1.0000   0.0162
  -8.250  -0.4681   0.07847   0.07627  -0.0387   1.0000   0.0162
  -8.000  -0.4735   0.07511   0.07288  -0.0389   1.0000   0.0163
  -7.750  -0.4767   0.07166   0.06938  -0.0389   1.0000   0.0163
  -7.500  -0.4785   0.06835   0.06602  -0.0382   1.0000   0.0163
  -7.250  -0.4796   0.06507   0.06266  -0.0371   1.0000   0.0163
  -7.000  -0.4799   0.06196   0.05946  -0.0355   1.0000   0.0163
  -6.750  -0.4930   0.05475   0.05214  -0.0342   0.9993   0.0168
  -6.500  -0.4714   0.04966   0.04692  -0.0382   0.9952   0.0174
  -6.250  -0.4463   0.04617   0.04332  -0.0412   0.9903   0.0180
  -6.000  -0.4182   0.04280   0.03980  -0.0442   0.9852   0.0187
  -5.750  -0.3896   0.03969   0.03650  -0.0466   0.9795   0.0199
  -5.500  -0.3597   0.03655   0.03314  -0.0484   0.9728   0.0217
  -5.250  -0.3186   0.03569   0.03196  -0.0497   0.9689   0.0242
  -5.000  -0.2893   0.03371   0.02968  -0.0504   0.9587   0.0245
  -4.750  -0.2696   0.02694   0.02248  -0.0514   0.9469   0.0260
  -4.500  -0.2404   0.02483   0.02028  -0.0529   0.9362   0.0271
  -4.250  -0.2109   0.02317   0.01846  -0.0539   0.9236   0.0286
  -4.000  -0.1822   0.02182   0.01690  -0.0543   0.9098   0.0311
  -3.750  -0.1501   0.02255   0.01738  -0.0542   0.8967   0.0351
  -3.500  -0.1284   0.01940   0.01379  -0.0532   0.8841   0.0369
  -3.250  -0.1039   0.01767   0.01198  -0.0531   0.8734   0.0389
  -3.000  -0.0788   0.01672   0.01092  -0.0527   0.8635   0.0413
  -2.750  -0.0524   0.01626   0.01029  -0.0522   0.8551   0.0460
  -2.500  -0.0275   0.01494   0.00878  -0.0516   0.8471   0.0520
  -2.250  -0.0014   0.01419   0.00797  -0.0513   0.8401   0.0570
  -1.750   0.0561   0.01243   0.00583  -0.0497   0.8271   0.0362
  -1.500   0.0817   0.01150   0.00491  -0.0492   0.8202   0.0342
  -1.250   0.1068   0.01085   0.00423  -0.0485   0.8141   0.0335
  -1.000   0.1305   0.01023   0.00358  -0.0476   0.8066   0.0317
  -0.750   0.1549   0.00982   0.00310  -0.0468   0.7997   0.0303
  -0.500   0.1793   0.00948   0.00271  -0.0460   0.7915   0.0293
  -0.250   0.2036   0.00922   0.00238  -0.0451   0.7794   0.0288
   0.000   0.2278   0.00900   0.00208  -0.0441   0.7635   0.0288
   0.250   0.2529   0.00883   0.00185  -0.0435   0.7502   0.0292
   0.500   0.2778   0.00870   0.00163  -0.0427   0.7328   0.0304
   0.750   0.3032   0.00861   0.00145  -0.0420   0.7160   0.0344
   1.000   0.3130   0.00675   0.00135  -0.0390   0.7023   0.6656
   1.250   0.3283   0.00608   0.00144  -0.0357   0.6830   0.8810
   1.500   0.3960   0.00632   0.00156  -0.0439   0.6296   0.9737
   1.750   0.4514   0.00689   0.00163  -0.0501   0.5192   0.9927
   2.000   0.4822   0.00792   0.00185  -0.0516   0.3384   0.9984
   2.250   0.5048   0.00860   0.00206  -0.0512   0.2360   1.0000
   2.500   0.5238   0.00894   0.00218  -0.0496   0.1908   1.0000
   2.750   0.5431   0.00927   0.00233  -0.0482   0.1510   1.0000
   3.250   0.5800   0.01019   0.00279  -0.0449   0.0511   1.0000
   3.500   0.6019   0.01038   0.00295  -0.0438   0.0482   1.0000
   3.750   0.6243   0.01056   0.00312  -0.0428   0.0463   1.0000
   4.000   0.6466   0.01076   0.00332  -0.0417   0.0438   1.0000
   4.250   0.6688   0.01099   0.00359  -0.0407   0.0416   1.0000
   4.500   0.6909   0.01122   0.00386  -0.0396   0.0406   1.0000
   4.750   0.7125   0.01152   0.00421  -0.0385   0.0386   1.0000
   5.000   0.7332   0.01191   0.00462  -0.0373   0.0354   1.0000
   5.250   0.7540   0.01230   0.00506  -0.0360   0.0329   1.0000
   5.500   0.7763   0.01255   0.00535  -0.0351   0.0313   1.0000
   5.750   0.7989   0.01277   0.00561  -0.0342   0.0294   1.0000
   6.000   0.8211   0.01306   0.00593  -0.0332   0.0267   1.0000
   6.250   0.8405   0.01360   0.00650  -0.0318   0.0230   1.0000
   6.500   0.8673   0.01347   0.00639  -0.0316   0.0179   1.0000
   6.750   0.8901   0.01371   0.00655  -0.0308   0.0132   1.0000
   7.000   0.9103   0.01422   0.00705  -0.0294   0.0113   1.0000
   7.250   0.9295   0.01484   0.00776  -0.0279   0.0108   1.0000
   7.500   0.9475   0.01557   0.00859  -0.0262   0.0103   1.0000
   7.750   0.9644   0.01639   0.00952  -0.0243   0.0101   1.0000
   8.000   0.9799   0.01735   0.01059  -0.0222   0.0099   1.0000
   8.250   0.9941   0.01846   0.01185  -0.0199   0.0098   1.0000
   8.500   1.0083   0.01966   0.01318  -0.0177   0.0098   1.0000
   8.750   1.0235   0.02079   0.01441  -0.0159   0.0094   1.0000
   9.000   1.0385   0.02210   0.01584  -0.0140   0.0091   1.0000
   9.250   1.0532   0.02357   0.01744  -0.0121   0.0088   1.0000
   9.500   1.0673   0.02540   0.01945  -0.0103   0.0086   1.0000
   9.750   1.0817   0.02774   0.02207  -0.0085   0.0089   1.0000
  15.500   0.6485   0.16535   0.16328  -0.0405   0.0194   1.0000
  15.750   0.6514   0.16860   0.16653  -0.0424   0.0183   1.0000
<< Back to BOEING 707 .99 SPAN AIRFOIL (b707e-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 707 .99 SPAN AIRFOIL (b707e-il)