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

NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA 63-210 AIRFOIL (n63210-il)
Reynolds number: 100,000
Max Cl/Cd: 42.8 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n63210-il-100000-n5.txt
Download as CSV file: xf-n63210-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-210 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6103   0.07608   0.07105  -0.0332   1.0000   0.0264
  -9.750  -0.6237   0.07035   0.06528  -0.0375   1.0000   0.0262
  -9.500  -0.6380   0.06559   0.06046  -0.0403   1.0000   0.0261
  -9.250  -0.6543   0.06153   0.05630  -0.0414   1.0000   0.0260
  -9.000  -0.6660   0.05721   0.05181  -0.0420   1.0000   0.0260
  -8.750  -0.6732   0.05282   0.04711  -0.0420   1.0000   0.0261
  -8.500  -0.6763   0.04848   0.04229  -0.0416   1.0000   0.0264
  -8.250  -0.6742   0.04416   0.03764  -0.0410   1.0000   0.0269
  -8.000  -0.6644   0.04127   0.03457  -0.0405   1.0000   0.0280
  -7.750  -0.6494   0.03966   0.03286  -0.0400   1.0000   0.0301
  -7.500  -0.6362   0.03709   0.02990  -0.0391   1.0000   0.0325
  -7.250  -0.6216   0.03428   0.02657  -0.0379   1.0000   0.0343
  -7.000  -0.6056   0.03193   0.02376  -0.0368   1.0000   0.0365
  -6.750  -0.5877   0.02999   0.02167  -0.0359   1.0000   0.0380
  -6.500  -0.5687   0.02862   0.02016  -0.0350   1.0000   0.0401
  -6.250  -0.5490   0.02704   0.01829  -0.0339   1.0000   0.0412
  -6.000  -0.5288   0.02545   0.01649  -0.0326   1.0000   0.0413
  -5.750  -0.5088   0.02409   0.01497  -0.0314   1.0000   0.0415
  -5.500  -0.4893   0.02292   0.01367  -0.0300   1.0000   0.0418
  -5.250  -0.4705   0.02188   0.01256  -0.0286   1.0000   0.0422
  -5.000  -0.4524   0.02098   0.01159  -0.0271   1.0000   0.0427
  -4.750  -0.4247   0.02005   0.01056  -0.0276   0.9958   0.0434
  -4.500  -0.3898   0.01916   0.00959  -0.0294   0.9883   0.0444
  -4.250  -0.3546   0.01840   0.00871  -0.0314   0.9809   0.0458
  -4.000  -0.3180   0.01765   0.00786  -0.0336   0.9744   0.0479
  -3.750  -0.2835   0.01700   0.00711  -0.0354   0.9662   0.0522
  -3.500  -0.2480   0.01644   0.00648  -0.0373   0.9586   0.0609
  -3.250  -0.2143   0.01543   0.00583  -0.0392   0.9508   0.1259
  -3.000  -0.1898   0.01353   0.00549  -0.0401   0.9407   0.4460
  -2.750  -0.1639   0.01321   0.00582  -0.0392   0.9311   0.6221
  -2.500  -0.1355   0.01324   0.00598  -0.0385   0.9218   0.6741
  -2.250  -0.1077   0.01326   0.00597  -0.0381   0.9111   0.7034
  -2.000  -0.0812   0.01335   0.00605  -0.0370   0.9014   0.7353
  -1.750  -0.0514   0.01334   0.00598  -0.0368   0.8932   0.7537
  -1.500  -0.0226   0.01329   0.00585  -0.0368   0.8833   0.7653
  -1.250   0.0072   0.01324   0.00571  -0.0370   0.8745   0.7770
  -1.000   0.0365   0.01320   0.00560  -0.0369   0.8663   0.7876
  -0.750   0.0636   0.01318   0.00554  -0.0365   0.8565   0.7984
  -0.500   0.0912   0.01317   0.00549  -0.0362   0.8474   0.8097
  -0.250   0.1187   0.01317   0.00544  -0.0359   0.8384   0.8217
   0.000   0.1443   0.01318   0.00545  -0.0353   0.8280   0.8339
   0.250   0.1701   0.01319   0.00546  -0.0346   0.8184   0.8458
   0.500   0.1963   0.01319   0.00546  -0.0339   0.8095   0.8582
   0.750   0.2214   0.01320   0.00551  -0.0332   0.7990   0.8713
   1.000   0.2474   0.01321   0.00554  -0.0325   0.7896   0.8848
   1.250   0.2744   0.01319   0.00554  -0.0321   0.7809   0.8988
   1.500   0.3017   0.01320   0.00562  -0.0318   0.7706   0.9135
   1.750   0.3312   0.01320   0.00566  -0.0320   0.7613   0.9286
   2.000   0.3632   0.01319   0.00570  -0.0327   0.7521   0.9434
   2.250   0.3976   0.01321   0.00582  -0.0340   0.7414   0.9584
   2.500   0.4342   0.01322   0.00591  -0.0359   0.7302   0.9738
   2.750   0.4708   0.01306   0.00576  -0.0372   0.7079   0.9928
   3.000   0.4944   0.01302   0.00572  -0.0362   0.6809   1.0000
   3.250   0.5171   0.01307   0.00576  -0.0351   0.6568   1.0000
   3.500   0.5404   0.01312   0.00578  -0.0340   0.6260   1.0000
   3.750   0.5621   0.01321   0.00569  -0.0325   0.5700   1.0000
   4.000   0.5817   0.01359   0.00564  -0.0307   0.4697   1.0000
   4.250   0.5950   0.01496   0.00591  -0.0287   0.2747   1.0000
   4.500   0.6105   0.01665   0.00674  -0.0276   0.1202   1.0000
   4.750   0.6304   0.01785   0.00757  -0.0268   0.0647   1.0000
   5.000   0.6532   0.01867   0.00839  -0.0262   0.0538   1.0000
   5.250   0.6754   0.01955   0.00932  -0.0256   0.0492   1.0000
   5.500   0.6981   0.02033   0.01023  -0.0249   0.0467   1.0000
   5.750   0.7203   0.02115   0.01116  -0.0242   0.0442   1.0000
   6.000   0.7421   0.02202   0.01208  -0.0236   0.0414   1.0000
   6.250   0.7628   0.02308   0.01316  -0.0229   0.0393   1.0000
   6.500   0.7833   0.02435   0.01448  -0.0220   0.0382   1.0000
   6.750   0.8060   0.02551   0.01574  -0.0214   0.0375   1.0000
   7.000   0.8295   0.02682   0.01715  -0.0208   0.0369   1.0000
   7.250   0.8537   0.02827   0.01873  -0.0204   0.0363   1.0000
   7.500   0.8783   0.02990   0.02052  -0.0200   0.0358   1.0000
   7.750   0.9026   0.03172   0.02255  -0.0195   0.0354   1.0000
   8.000   0.9260   0.03378   0.02491  -0.0190   0.0351   1.0000
   8.250   0.9475   0.03608   0.02753  -0.0183   0.0348   1.0000
   8.500   0.9666   0.03867   0.03051  -0.0175   0.0347   1.0000
   8.750   0.9826   0.04139   0.03364  -0.0164   0.0342   1.0000
   9.000   0.9966   0.04371   0.03628  -0.0154   0.0326   1.0000
   9.250   1.0105   0.04576   0.03851  -0.0146   0.0310   1.0000
   9.500   1.0216   0.04820   0.04107  -0.0139   0.0296   1.0000
   9.750   1.0225   0.05176   0.04502  -0.0123   0.0285   1.0000
  10.000   1.0183   0.05532   0.04910  -0.0103   0.0279   1.0000
  10.250   1.0085   0.05917   0.05334  -0.0084   0.0276   1.0000
  10.500   0.9926   0.06278   0.05723  -0.0063   0.0275   1.0000
  10.750   0.9745   0.06675   0.06144  -0.0055   0.0275   1.0000
  11.000   0.9543   0.07138   0.06623  -0.0061   0.0274   1.0000
  11.250   0.9338   0.07669   0.07172  -0.0082   0.0275   1.0000
  11.500   0.9132   0.08284   0.07802  -0.0116   0.0277   1.0000
  11.750   0.8917   0.09016   0.08546  -0.0164   0.0279   1.0000
<< Back to NACA 63-210 AIRFOIL (n63210-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-210 AIRFOIL (n63210-il)