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

ARA-D 10% AIRFOIL (arad10-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: ARA-D 10% AIRFOIL (arad10-il)
Reynolds number: 100,000
Max Cl/Cd: 51.27 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-arad10-il-100000.txt
Download as CSV file: xf-arad10-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: ARA-D 10% AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3582   0.10554   0.10046  -0.0207   1.0000   0.0941
  -8.250  -0.3750   0.10421   0.09924  -0.0250   1.0000   0.0961
  -8.000  -0.3921   0.10242   0.09753  -0.0335   1.0000   0.0967
  -7.750  -0.3740   0.09676   0.09193  -0.0266   1.0000   0.0981
  -7.500  -0.3580   0.09344   0.08863  -0.0231   1.0000   0.1007
  -7.250  -0.3549   0.09081   0.08606  -0.0229   1.0000   0.1037
  -7.000  -0.3580   0.08829   0.08361  -0.0247   1.0000   0.1074
  -6.750  -0.3699   0.08608   0.08131  -0.0394   1.0000   0.1107
  -6.500  -0.3647   0.08162   0.07702  -0.0324   1.0000   0.1120
  -6.250  -0.3596   0.07898   0.07445  -0.0282   1.0000   0.1137
  -6.000  -0.3558   0.07655   0.07208  -0.0265   1.0000   0.1161
  -5.750  -0.3518   0.07397   0.06951  -0.0268   1.0000   0.1197
  -5.500  -0.3390   0.06971   0.06501  -0.0367   1.0000   0.1267
  -5.250  -0.3351   0.06686   0.06228  -0.0332   1.0000   0.1281
  -5.000  -0.3288   0.06455   0.06003  -0.0310   1.0000   0.1305
  -4.750  -0.3021   0.06163   0.05668  -0.0394   1.0000   0.1424
  -4.500  -0.2972   0.05824   0.05348  -0.0365   1.0000   0.1441
  -4.250  -0.2893   0.05622   0.05155  -0.0341   1.0000   0.1478
  -4.000  -0.2260   0.04194   0.03565  -0.0486   1.0000   0.0967
  -3.750  -0.2074   0.03899   0.03273  -0.0486   1.0000   0.0949
  -3.500  -0.1772   0.03505   0.02818  -0.0506   1.0000   0.0954
  -3.250  -0.1332   0.03188   0.02461  -0.0547   0.9962   0.0997
  -3.000  -0.0855   0.02994   0.02222  -0.0588   0.9908   0.1085
  -2.750  -0.0390   0.02841   0.02051  -0.0630   0.9850   0.1184
  -2.500   0.0075   0.02684   0.01852  -0.0667   0.9786   0.1294
  -2.250   0.0561   0.02596   0.01731  -0.0706   0.9721   0.1421
  -2.000   0.0988   0.02498   0.01633  -0.0738   0.9637   0.1551
  -1.750   0.1475   0.02404   0.01530  -0.0779   0.9574   0.1706
  -1.500   0.1890   0.02328   0.01452  -0.0806   0.9477   0.1871
  -1.250   0.2399   0.02248   0.01364  -0.0848   0.9416   0.2099
  -1.000   0.2809   0.02170   0.01305  -0.0874   0.9310   0.2332
  -0.750   0.3367   0.02054   0.01206  -0.0923   0.9248   0.2689
  -0.500   0.3782   0.01955   0.01128  -0.0941   0.9112   0.3113
  -0.250   0.4179   0.01818   0.01059  -0.0956   0.8981   0.4300
   0.000   0.4500   0.01644   0.00995  -0.0942   0.8847   1.0000
   0.250   0.4841   0.01619   0.00946  -0.0942   0.8705   1.0000
   0.500   0.5149   0.01598   0.00908  -0.0936   0.8553   1.0000
   0.750   0.5433   0.01583   0.00878  -0.0926   0.8391   1.0000
   1.000   0.5703   0.01573   0.00856  -0.0914   0.8216   1.0000
   1.250   0.5966   0.01567   0.00838  -0.0901   0.8033   1.0000
   1.500   0.6228   0.01564   0.00824  -0.0889   0.7843   1.0000
   1.750   0.6490   0.01562   0.00812  -0.0876   0.7644   1.0000
   2.000   0.6752   0.01562   0.00799  -0.0864   0.7439   1.0000
   2.250   0.7016   0.01564   0.00787  -0.0852   0.7230   1.0000
   2.500   0.7272   0.01573   0.00786  -0.0840   0.6983   1.0000
   2.750   0.7532   0.01581   0.00779  -0.0829   0.6733   1.0000
   3.000   0.7787   0.01594   0.00781  -0.0818   0.6445   1.0000
   3.250   0.8041   0.01611   0.00783  -0.0806   0.6130   1.0000
   3.500   0.8290   0.01634   0.00787  -0.0795   0.5774   1.0000
   3.750   0.8534   0.01666   0.00794  -0.0783   0.5380   1.0000
   4.000   0.8773   0.01711   0.00815  -0.0772   0.4938   1.0000
   4.250   0.9008   0.01770   0.00840  -0.0761   0.4533   1.0000
   4.500   0.9246   0.01838   0.00880  -0.0752   0.4166   1.0000
   4.750   0.9487   0.01910   0.00926  -0.0745   0.3877   1.0000
   5.000   0.9733   0.01982   0.00982  -0.0739   0.3631   1.0000
   5.250   0.9982   0.02057   0.01039  -0.0734   0.3435   1.0000
   5.500   1.0233   0.02133   0.01103  -0.0730   0.3268   1.0000
   5.750   1.0486   0.02212   0.01171  -0.0726   0.3123   1.0000
   6.000   1.0740   0.02294   0.01244  -0.0722   0.2997   1.0000
   6.250   1.0999   0.02382   0.01314  -0.0720   0.2885   1.0000
   6.500   1.1248   0.02458   0.01398  -0.0715   0.2773   1.0000
   6.750   1.1503   0.02550   0.01487  -0.0712   0.2678   1.0000
   7.000   1.1757   0.02635   0.01566  -0.0710   0.2586   1.0000
   7.250   1.2003   0.02732   0.01674  -0.0705   0.2500   1.0000
   7.500   1.2259   0.02825   0.01759  -0.0703   0.2422   1.0000
   7.750   1.2497   0.02930   0.01878  -0.0698   0.2345   1.0000
   8.000   1.2742   0.03023   0.01971  -0.0694   0.2272   1.0000
   8.250   1.2982   0.03145   0.02100  -0.0690   0.2207   1.0000
   8.500   1.3205   0.03245   0.02217  -0.0683   0.2138   1.0000
   8.750   1.3465   0.03375   0.02332  -0.0683   0.2079   1.0000
   9.000   1.3652   0.03492   0.02486  -0.0671   0.2016   1.0000
   9.250   1.3879   0.03599   0.02598  -0.0665   0.1959   1.0000
   9.500   1.4099   0.03752   0.02758  -0.0660   0.1908   1.0000
   9.750   1.4270   0.03890   0.02928  -0.0646   0.1851   1.0000
  10.000   1.4485   0.04007   0.03048  -0.0640   0.1802   1.0000
  10.250   1.4680   0.04186   0.03237  -0.0632   0.1758   1.0000
  10.500   1.4794   0.04362   0.03454  -0.0612   0.1709   1.0000
  10.750   1.4977   0.04495   0.03597  -0.0602   0.1664   1.0000
  11.000   1.5223   0.04672   0.03759  -0.0603   0.1624   1.0000
  11.250   1.5207   0.04912   0.04059  -0.0569   0.1588   1.0000
  11.500   1.5265   0.05125   0.04304  -0.0545   0.1549   1.0000
  11.750   1.5435   0.05261   0.04445  -0.0535   0.1512   1.0000
  12.000   1.5641   0.05467   0.04643  -0.0533   0.1477   1.0000
  12.250   1.5453   0.05801   0.05034  -0.0484   0.1456   1.0000
  12.500   1.5243   0.06157   0.05431  -0.0439   0.1436   1.0000
  12.750   1.4975   0.06514   0.05819  -0.0390   0.1421   1.0000
  13.000   1.4638   0.06961   0.06295  -0.0350   0.1412   1.0000
  13.250   1.4051   0.07732   0.07101  -0.0329   0.1416   1.0000
  13.500   1.2708   0.09858   0.09266  -0.0402   0.1457   1.0000
  13.750   1.4391   0.07959   0.07326  -0.0313   0.1356   1.0000
  14.000   1.4086   0.08597   0.07981  -0.0315   0.1348   1.0000
  14.250   1.0276   0.18600   0.18016  -0.0880   0.1926   1.0000
<< Back to ARA-D 10% AIRFOIL (arad10-il)

Polar data table (+)

Polar graphs


<< Back to ARA-D 10% AIRFOIL (arad10-il)