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

NACA 63-015A AIRFOIL (n63015a-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 63-015A AIRFOIL (n63015a-il)
Reynolds number: 50,000
Max Cl/Cd: 28.49 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n63015a-il-50000-n5.txt
Download as CSV file: xf-n63015a-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-015A AIRFOIL                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.000  -0.7540   0.10804   0.10026  -0.0228   1.0000   0.0564
 -13.750  -0.7819   0.09892   0.09103  -0.0281   1.0000   0.0561
 -13.500  -0.8098   0.09088   0.08283  -0.0326   1.0000   0.0558
 -13.250  -0.8356   0.08396   0.07572  -0.0360   1.0000   0.0557
 -13.000  -0.8586   0.07801   0.06954  -0.0383   1.0000   0.0556
 -12.750  -0.8783   0.07286   0.06414  -0.0396   1.0000   0.0557
 -12.500  -0.8946   0.06833   0.05933  -0.0401   1.0000   0.0559
 -12.250  -0.9072   0.06432   0.05500  -0.0399   1.0000   0.0562
 -12.000  -0.9147   0.06077   0.05116  -0.0392   1.0000   0.0568
 -11.750  -0.9067   0.05824   0.04859  -0.0386   1.0000   0.0582
 -11.500  -0.9013   0.05587   0.04611  -0.0378   1.0000   0.0599
 -11.250  -0.8964   0.05345   0.04350  -0.0368   1.0000   0.0619
 -11.000  -0.8892   0.05099   0.04075  -0.0357   1.0000   0.0642
 -10.750  -0.8777   0.04856   0.03792  -0.0347   1.0000   0.0666
 -10.500  -0.8558   0.04667   0.03609  -0.0343   1.0000   0.0697
 -10.250  -0.8393   0.04497   0.03428  -0.0335   1.0000   0.0739
 -10.000  -0.8177   0.04325   0.03232  -0.0328   1.0000   0.0791
  -9.750  -0.7985   0.04179   0.03091  -0.0320   1.0000   0.0845
  -9.500  -0.7789   0.04036   0.02932  -0.0311   1.0000   0.0920
  -9.250  -0.7646   0.03891   0.02794  -0.0299   1.0000   0.0994
  -9.000  -0.7521   0.03745   0.02647  -0.0284   1.0000   0.1085
  -8.750  -0.7416   0.03599   0.02499  -0.0268   1.0000   0.1196
  -8.500  -0.7344   0.03451   0.02357  -0.0248   1.0000   0.1328
  -8.250  -0.7304   0.03301   0.02220  -0.0224   1.0000   0.1478
  -8.000  -0.7291   0.03151   0.02089  -0.0195   1.0000   0.1660
  -7.750  -0.7273   0.03003   0.01960  -0.0166   1.0000   0.1915
  -7.500  -0.7279   0.02855   0.01844  -0.0133   1.0000   0.2249
  -7.250  -0.7296   0.02718   0.01749  -0.0095   1.0000   0.2702
  -7.000  -0.7305   0.02614   0.01692  -0.0053   1.0000   0.3285
  -6.750  -0.7272   0.02559   0.01675  -0.0012   1.0000   0.3902
  -6.500  -0.7216   0.02533   0.01663   0.0027   1.0000   0.4398
  -6.250  -0.7134   0.02529   0.01666   0.0064   1.0000   0.4782
  -6.000  -0.7047   0.02531   0.01666   0.0100   1.0000   0.5096
  -5.750  -0.6968   0.02529   0.01656   0.0135   1.0000   0.5367
  -5.500  -0.6846   0.02553   0.01677   0.0167   1.0000   0.5600
  -5.250  -0.6755   0.02559   0.01675   0.0200   1.0000   0.5826
  -5.000  -0.6624   0.02587   0.01697   0.0230   1.0000   0.6027
  -4.750  -0.6457   0.02618   0.01721   0.0254   0.9990   0.6218
  -4.500  -0.6081   0.02662   0.01749   0.0239   0.9897   0.6425
  -4.250  -0.5709   0.02689   0.01760   0.0223   0.9804   0.6600
  -4.000  -0.5323   0.02709   0.01762   0.0204   0.9717   0.6744
  -3.750  -0.4958   0.02713   0.01751   0.0186   0.9621   0.6871
  -3.500  -0.4619   0.02700   0.01721   0.0170   0.9520   0.6996
  -3.250  -0.4207   0.02699   0.01705   0.0143   0.9446   0.7100
  -3.000  -0.3870   0.02691   0.01685   0.0129   0.9343   0.7192
  -2.750  -0.3526   0.02665   0.01644   0.0109   0.9253   0.7305
  -2.500  -0.3139   0.02669   0.01639   0.0089   0.9168   0.7378
  -2.250  -0.2829   0.02646   0.01606   0.0077   0.9071   0.7481
  -2.000  -0.2449   0.02645   0.01596   0.0058   0.8988   0.7556
  -1.750  -0.2167   0.02627   0.01571   0.0052   0.8887   0.7655
  -1.500  -0.1787   0.02623   0.01560   0.0033   0.8810   0.7731
  -1.250  -0.1528   0.02611   0.01544   0.0033   0.8705   0.7827
  -1.000  -0.1151   0.02605   0.01533   0.0015   0.8634   0.7905
  -0.750  -0.0907   0.02600   0.01525   0.0017   0.8524   0.7999
  -0.500  -0.0561   0.02595   0.01516   0.0005   0.8449   0.8081
  -0.250  -0.0298   0.02594   0.01515   0.0005   0.8347   0.8171
   0.000   0.0000   0.02592   0.01512   0.0000   0.8262   0.8262
   0.250   0.0298   0.02594   0.01515  -0.0005   0.8171   0.8347
   0.500   0.0561   0.02595   0.01516  -0.0005   0.8081   0.8450
   0.750   0.0907   0.02600   0.01525  -0.0017   0.7998   0.8524
   1.000   0.1151   0.02605   0.01533  -0.0015   0.7905   0.8634
   1.250   0.1528   0.02611   0.01544  -0.0033   0.7827   0.8705
   1.500   0.1788   0.02623   0.01560  -0.0033   0.7731   0.8810
   1.750   0.2167   0.02627   0.01571  -0.0052   0.7655   0.8888
   2.000   0.2449   0.02644   0.01596  -0.0058   0.7557   0.8989
   2.250   0.2829   0.02646   0.01606  -0.0077   0.7481   0.9072
   2.500   0.3139   0.02669   0.01639  -0.0089   0.7378   0.9168
   2.750   0.3526   0.02665   0.01644  -0.0109   0.7305   0.9254
   3.000   0.3871   0.02691   0.01685  -0.0129   0.7192   0.9343
   3.250   0.4207   0.02699   0.01704  -0.0143   0.7100   0.9446
   3.500   0.4619   0.02700   0.01721  -0.0170   0.6996   0.9521
   3.750   0.4958   0.02712   0.01750  -0.0186   0.6871   0.9621
   4.000   0.5323   0.02708   0.01762  -0.0204   0.6744   0.9717
   4.250   0.5709   0.02689   0.01760  -0.0223   0.6600   0.9804
   4.500   0.6082   0.02662   0.01749  -0.0239   0.6425   0.9897
   4.750   0.6458   0.02618   0.01720  -0.0254   0.6218   0.9990
   5.000   0.6623   0.02586   0.01697  -0.0230   0.6027   1.0000
   5.250   0.6754   0.02559   0.01674  -0.0200   0.5827   1.0000
   5.500   0.6845   0.02553   0.01677  -0.0167   0.5601   1.0000
   5.750   0.6968   0.02529   0.01656  -0.0135   0.5367   1.0000
   6.000   0.7046   0.02530   0.01665  -0.0100   0.5096   1.0000
   6.250   0.7133   0.02529   0.01665  -0.0064   0.4782   1.0000
   6.500   0.7216   0.02533   0.01663  -0.0027   0.4398   1.0000
   6.750   0.7272   0.02559   0.01674   0.0012   0.3902   1.0000
   7.000   0.7305   0.02613   0.01692   0.0053   0.3285   1.0000
   7.250   0.7296   0.02718   0.01748   0.0095   0.2701   1.0000
   7.500   0.7279   0.02855   0.01844   0.0133   0.2248   1.0000
   7.750   0.7274   0.03003   0.01960   0.0166   0.1914   1.0000
   8.000   0.7292   0.03151   0.02089   0.0195   0.1660   1.0000
   8.250   0.7306   0.03301   0.02220   0.0224   0.1478   1.0000
   8.500   0.7346   0.03451   0.02357   0.0248   0.1328   1.0000
   8.750   0.7419   0.03599   0.02499   0.0267   0.1196   1.0000
   9.000   0.7525   0.03744   0.02647   0.0284   0.1085   1.0000
   9.250   0.7651   0.03891   0.02793   0.0298   0.0993   1.0000
   9.500   0.7794   0.04036   0.02932   0.0310   0.0919   1.0000
   9.750   0.7990   0.04179   0.03091   0.0319   0.0844   1.0000
  10.000   0.8184   0.04325   0.03231   0.0327   0.0790   1.0000
  10.250   0.8399   0.04497   0.03428   0.0334   0.0739   1.0000
  10.500   0.8564   0.04667   0.03609   0.0342   0.0697   1.0000
  10.750   0.8783   0.04857   0.03793   0.0346   0.0666   1.0000
  11.000   0.8898   0.05100   0.04076   0.0357   0.0642   1.0000
  11.250   0.8970   0.05346   0.04351   0.0367   0.0619   1.0000
  11.500   0.9020   0.05588   0.04612   0.0377   0.0598   1.0000
  11.750   0.9075   0.05825   0.04859   0.0385   0.0581   1.0000
  12.000   0.9156   0.06078   0.05117   0.0391   0.0568   1.0000
  12.250   0.9078   0.06435   0.05503   0.0398   0.0562   1.0000
  12.500   0.8952   0.06838   0.05937   0.0399   0.0559   1.0000
  12.750   0.8790   0.07291   0.06419   0.0394   0.0557   1.0000
  13.000   0.8593   0.07807   0.06961   0.0381   0.0556   1.0000
  13.250   0.8362   0.08404   0.07581   0.0358   0.0556   1.0000
  13.500   0.8104   0.09099   0.08294   0.0324   0.0558   1.0000
  13.750   0.7826   0.09905   0.09116   0.0279   0.0561   1.0000
  14.000   0.7546   0.10823   0.10044   0.0225   0.0564   1.0000
<< Back to NACA 63-015A AIRFOIL (n63015a-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-015A AIRFOIL (n63015a-il)