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 .08 SPAN AIRFOIL (b707a-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: BOEING 707 .08 SPAN AIRFOIL (b707a-il)
Reynolds number: 500,000
Max Cl/Cd: 59.12 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-b707a-il-500000.txt
Download as CSV file: xf-b707a-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 707 .08 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
  ------ -------- --------- --------- -------- -------- --------
 -16.000  -0.8487   0.13023   0.12753   0.0085   1.0000   0.0215
 -15.750  -0.9599   0.09874   0.09568  -0.0105   1.0000   0.0205
 -15.500  -1.0391   0.08064   0.07711  -0.0213   1.0000   0.0198
 -15.250  -1.0854   0.07018   0.06627  -0.0271   1.0000   0.0194
 -15.000  -1.1197   0.06223   0.05794  -0.0311   1.0000   0.0192
 -14.750  -1.1366   0.05699   0.05243  -0.0331   1.0000   0.0190
 -14.500  -1.1466   0.05291   0.04813  -0.0345   1.0000   0.0188
 -14.250  -1.1536   0.04940   0.04441  -0.0354   1.0000   0.0187
 -14.000  -1.1558   0.04657   0.04143  -0.0358   1.0000   0.0187
 -13.750  -1.1571   0.04406   0.03875  -0.0359   1.0000   0.0186
 -13.500  -1.1558   0.04190   0.03645  -0.0356   1.0000   0.0186
 -13.250  -1.1533   0.03990   0.03432  -0.0350   1.0000   0.0185
 -13.000  -1.1493   0.03815   0.03246  -0.0341   1.0000   0.0185
 -12.750  -1.1428   0.03629   0.03049  -0.0330   1.0000   0.0184
 -12.500  -1.1358   0.03455   0.02864  -0.0318   1.0000   0.0184
 -12.250  -1.1267   0.03294   0.02694  -0.0304   1.0000   0.0184
 -12.000  -1.1188   0.03151   0.02540  -0.0288   1.0000   0.0184
 -11.750  -1.1102   0.03022   0.02402  -0.0271   1.0000   0.0184
 -11.500  -1.1017   0.02907   0.02278  -0.0251   1.0000   0.0184
 -11.250  -1.0922   0.02803   0.02166  -0.0230   1.0000   0.0185
 -11.000  -1.0824   0.02685   0.02042  -0.0207   1.0000   0.0185
 -10.750  -1.0730   0.02595   0.01947  -0.0182   1.0000   0.0185
 -10.500  -1.0636   0.02502   0.01849  -0.0155   1.0000   0.0184
 -10.250  -1.0524   0.02407   0.01750  -0.0131   1.0000   0.0184
 -10.000  -1.0403   0.02315   0.01654  -0.0108   1.0000   0.0184
  -9.750  -1.0279   0.02229   0.01564  -0.0085   1.0000   0.0184
  -9.500  -1.0150   0.02147   0.01479  -0.0062   1.0000   0.0184
  -9.250  -1.0018   0.02068   0.01397  -0.0039   1.0000   0.0185
  -9.000  -0.9881   0.01992   0.01320  -0.0016   1.0000   0.0185
  -8.750  -0.9750   0.01912   0.01237   0.0008   1.0000   0.0185
  -8.500  -0.9611   0.01835   0.01159   0.0031   1.0000   0.0186
  -8.250  -0.9468   0.01761   0.01082   0.0054   1.0000   0.0187
  -8.000  -0.9309   0.01696   0.01015   0.0074   1.0000   0.0188
  -7.750  -0.9145   0.01632   0.00949   0.0094   1.0000   0.0190
  -7.500  -0.8795   0.01559   0.00874   0.0075   0.9960   0.0193
  -7.250  -0.8484   0.01499   0.00812   0.0066   0.9915   0.0196
  -7.000  -0.8186   0.01446   0.00756   0.0060   0.9862   0.0201
  -6.750  -0.7863   0.01398   0.00705   0.0049   0.9812   0.0205
  -6.500  -0.7591   0.01356   0.00660   0.0050   0.9740   0.0210
  -6.250  -0.7263   0.01306   0.00609   0.0039   0.9668   0.0224
  -6.000  -0.6956   0.01259   0.00563   0.0033   0.9565   0.0245
  -5.750  -0.6391   0.01282   0.00521  -0.0026   0.6668   0.0415
  -5.500  -0.6186   0.01316   0.00512  -0.0010   0.5309   0.0464
  -5.250  -0.5932   0.01318   0.00502  -0.0004   0.4938   0.0501
  -5.000  -0.5694   0.01328   0.00485   0.0004   0.3886   0.0531
  -4.750  -0.5437   0.01324   0.00472   0.0009   0.3726   0.0560
  -4.500  -0.5178   0.01312   0.00456   0.0014   0.3672   0.0591
  -4.250  -0.4923   0.01297   0.00439   0.0020   0.3624   0.0622
  -4.000  -0.4661   0.01290   0.00427   0.0024   0.3588   0.0648
  -3.750  -0.4413   0.01262   0.00398   0.0031   0.3563   0.0679
  -3.500  -0.4156   0.01245   0.00378   0.0036   0.3547   0.0702
  -3.250  -0.3894   0.01230   0.00360   0.0041   0.3531   0.0721
  -3.000  -0.3634   0.01213   0.00340   0.0047   0.3511   0.0743
  -2.750  -0.3377   0.01193   0.00320   0.0052   0.3484   0.0781
  -2.500  -0.3115   0.01180   0.00305   0.0057   0.3457   0.0818
  -2.250  -0.2859   0.01161   0.00288   0.0063   0.3433   0.0921
  -2.000  -0.2686   0.01031   0.00231   0.0078   0.3407   0.2746
  -1.750  -0.2446   0.01011   0.00245   0.0086   0.3371   0.3902
  -1.500  -0.2171   0.01013   0.00249   0.0089   0.3338   0.4067
  -1.250  -0.1899   0.01017   0.00255   0.0093   0.3306   0.4185
  -1.000  -0.1626   0.01023   0.00259   0.0097   0.3276   0.4275
  -0.750  -0.1359   0.01032   0.00267   0.0101   0.3232   0.4341
  -0.500  -0.1085   0.01041   0.00273   0.0105   0.3190   0.4392
  -0.250  -0.0806   0.01047   0.00275   0.0107   0.3143   0.4432
   0.000  -0.0538   0.01047   0.00277   0.0111   0.3091   0.4482
   0.250  -0.0268   0.01055   0.00284   0.0115   0.3041   0.4522
   0.500   0.0010   0.01053   0.00285   0.0118   0.2988   0.4566
   0.750   0.0283   0.01062   0.00288   0.0121   0.2927   0.4607
   1.000   0.0564   0.01061   0.00289   0.0124   0.2861   0.4637
   1.250   0.0837   0.01053   0.00283   0.0127   0.2789   0.4673
   1.500   0.1120   0.01037   0.00276   0.0129   0.2680   0.4720
   2.250   0.1941   0.01051   0.00260   0.0138   0.1833   0.4847
   2.500   0.2207   0.01055   0.00267   0.0143   0.1737   0.4893
   2.750   0.2474   0.01063   0.00276   0.0147   0.1634   0.4933
   3.000   0.2741   0.01074   0.00284   0.0151   0.1523   0.4975
   3.250   0.3011   0.01083   0.00292   0.0155   0.1421   0.5019
   3.500   0.3277   0.01097   0.00301   0.0159   0.1291   0.5066
   3.750   0.3540   0.01106   0.00312   0.0163   0.1184   0.5134
   4.000   0.3803   0.01120   0.00325   0.0168   0.1071   0.5212
   4.250   0.4063   0.01135   0.00340   0.0173   0.0940   0.5312
   4.500   0.4316   0.01156   0.00356   0.0179   0.0734   0.5456
   4.750   0.4563   0.01174   0.00378   0.0186   0.0650   0.5701
   5.000   0.4799   0.01178   0.00404   0.0195   0.0606   0.6223
   5.250   0.5012   0.01175   0.00433   0.0209   0.0577   0.7147
   5.500   0.5219   0.01160   0.00456   0.0228   0.0558   0.8081
   5.750   0.5647   0.01169   0.00500   0.0200   0.0530   0.9331
   6.000   0.6196   0.01203   0.00537   0.0143   0.0501   0.9714
   6.250   0.6605   0.01235   0.00568   0.0116   0.0476   0.9868
   6.500   0.7005   0.01263   0.00597   0.0091   0.0457   0.9969
   6.750   0.7343   0.01286   0.00620   0.0079   0.0431   1.0000
   7.000   0.7572   0.01307   0.00642   0.0090   0.0407   1.0000
   7.250   0.7802   0.01329   0.00664   0.0101   0.0382   1.0000
   7.500   0.8029   0.01358   0.00686   0.0112   0.0257   1.0000
   7.750   0.8229   0.01419   0.00738   0.0127   0.0197   1.0000
   8.000   0.8442   0.01467   0.00789   0.0140   0.0186   1.0000
   8.250   0.8658   0.01512   0.00837   0.0153   0.0180   1.0000
   8.500   0.8875   0.01556   0.00886   0.0165   0.0177   1.0000
   8.750   0.9090   0.01601   0.00936   0.0177   0.0175   1.0000
   9.000   0.9304   0.01649   0.00991   0.0190   0.0173   1.0000
   9.250   0.9512   0.01703   0.01051   0.0203   0.0170   1.0000
   9.500   0.9719   0.01757   0.01111   0.0216   0.0169   1.0000
   9.750   0.9922   0.01815   0.01176   0.0229   0.0168   1.0000
  10.000   1.0122   0.01876   0.01245   0.0242   0.0167   1.0000
  10.250   1.0313   0.01944   0.01320   0.0256   0.0166   1.0000
  10.500   1.0499   0.02016   0.01400   0.0270   0.0165   1.0000
  10.750   1.0678   0.02093   0.01483   0.0285   0.0164   1.0000
  11.000   1.0847   0.02175   0.01573   0.0301   0.0164   1.0000
  11.250   1.1005   0.02263   0.01669   0.0317   0.0163   1.0000
  11.500   1.1148   0.02360   0.01772   0.0335   0.0162   1.0000
  11.750   1.1276   0.02462   0.01883   0.0353   0.0161   1.0000
  12.000   1.1385   0.02573   0.02001   0.0374   0.0160   1.0000
  12.250   1.1475   0.02689   0.02124   0.0395   0.0159   1.0000
  12.500   1.1536   0.02814   0.02257   0.0420   0.0159   1.0000
  12.750   1.1551   0.02935   0.02386   0.0451   0.0158   1.0000
  13.000   1.1532   0.03076   0.02536   0.0481   0.0158   1.0000
  13.250   1.1506   0.03241   0.02710   0.0505   0.0158   1.0000
  13.500   1.1461   0.03451   0.02931   0.0519   0.0158   1.0000
  13.750   1.1387   0.03746   0.03238   0.0516   0.0158   1.0000
  14.000   1.1297   0.04212   0.03720   0.0481   0.0158   1.0000
  14.250   1.1191   0.04825   0.04350   0.0429   0.0159   1.0000
  14.500   1.1095   0.05447   0.04987   0.0382   0.0159   1.0000
  14.750   1.0946   0.06070   0.05620   0.0342   0.0159   1.0000
  15.000   1.0785   0.06669   0.06229   0.0306   0.0160   1.0000
  15.250   1.0627   0.07240   0.06810   0.0273   0.0161   1.0000
  15.500   1.0481   0.07789   0.07367   0.0244   0.0161   1.0000
  15.750   1.0342   0.08351   0.07939   0.0213   0.0162   1.0000
  16.000   1.0229   0.08870   0.08465   0.0186   0.0162   1.0000
<< Back to BOEING 707 .08 SPAN AIRFOIL (b707a-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 707 .08 SPAN AIRFOIL (b707a-il)