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

NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA M21 AIRFOIL (m21-il)
Reynolds number: 500,000
Max Cl/Cd: 105.82 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m21-il-500000-n5.txt
Download as CSV file: xf-m21-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M21 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3130   0.09657   0.09312  -0.0178   0.6466   0.0200
  -8.750  -0.3102   0.09278   0.08933  -0.0200   0.6419   0.0205
  -8.500  -0.3259   0.08489   0.08145  -0.0265   0.6388   0.0217
  -8.250  -0.3199   0.08242   0.07894  -0.0284   0.6336   0.0219
  -8.000  -0.3106   0.08034   0.07684  -0.0297   0.6287   0.0222
  -7.750  -0.2997   0.07779   0.07427  -0.0318   0.6235   0.0224
  -7.500  -0.2890   0.07479   0.07121  -0.0343   0.6188   0.0227
  -7.250  -0.2775   0.07153   0.06789  -0.0369   0.6147   0.0231
  -7.000  -0.2649   0.06801   0.06433  -0.0396   0.6104   0.0234
  -6.750  -0.2513   0.06424   0.06048  -0.0422   0.6058   0.0239
  -5.250  -0.1846   0.01961   0.01329  -0.0496   0.5860   0.0307
  -5.000  -0.1611   0.01756   0.01085  -0.0492   0.5818   0.0313
  -4.750  -0.1343   0.01680   0.00995  -0.0491   0.5770   0.0317
  -4.500  -0.1070   0.01624   0.00927  -0.0490   0.5721   0.0321
  -4.250  -0.0796   0.01569   0.00858  -0.0489   0.5678   0.0326
  -4.000  -0.0520   0.01515   0.00791  -0.0489   0.5635   0.0331
  -3.750  -0.0242   0.01461   0.00725  -0.0489   0.5586   0.0337
  -3.500   0.0037   0.01412   0.00662  -0.0489   0.5540   0.0343
  -3.250   0.0316   0.01368   0.00605  -0.0489   0.5500   0.0350
  -3.000   0.0598   0.01328   0.00555  -0.0489   0.5456   0.0356
  -2.750   0.0879   0.01296   0.00513  -0.0489   0.5409   0.0362
  -2.500   0.1158   0.01266   0.00478  -0.0490   0.5365   0.0373
  -2.250   0.1441   0.01249   0.00459  -0.0490   0.5325   0.0382
  -2.000   0.1724   0.01231   0.00438  -0.0491   0.5281   0.0392
  -1.750   0.2005   0.01212   0.00415  -0.0491   0.5235   0.0403
  -1.500   0.2284   0.01195   0.00391  -0.0491   0.5194   0.0415
  -1.250   0.2566   0.01178   0.00369  -0.0491   0.5155   0.0426
  -1.000   0.2845   0.01161   0.00354  -0.0492   0.5110   0.0443
  -0.750   0.3127   0.01155   0.00346  -0.0492   0.5066   0.0460
  -0.250   0.3690   0.01142   0.00327  -0.0494   0.4987   0.0502
   0.000   0.3970   0.01134   0.00322  -0.0495   0.4943   0.0528
   0.250   0.4251   0.01133   0.00319  -0.0495   0.4899   0.0551
   0.500   0.4531   0.01131   0.00314  -0.0496   0.4861   0.0576
   0.750   0.4812   0.01125   0.00307  -0.0497   0.4823   0.0596
   1.000   0.5091   0.01121   0.00306  -0.0498   0.4779   0.0624
   1.250   0.5371   0.01123   0.00307  -0.0499   0.4735   0.0653
   1.500   0.5651   0.01125   0.00307  -0.0500   0.4696   0.0681
   1.750   0.5931   0.01122   0.00305  -0.0501   0.4657   0.0704
   2.000   0.6206   0.01117   0.00303  -0.0501   0.4615   0.0733
   2.250   0.6483   0.01119   0.00305  -0.0502   0.4573   0.0762
   2.500   0.6760   0.01121   0.00306  -0.0503   0.4533   0.0791
   2.750   0.7039   0.01122   0.00308  -0.0504   0.4491   0.0812
   3.000   0.7314   0.01122   0.00309  -0.0505   0.4450   0.0840
   3.250   0.7587   0.01125   0.00313  -0.0506   0.4410   0.0879
   3.500   0.7863   0.01129   0.00319  -0.0507   0.4370   0.0915
   3.750   0.8140   0.01133   0.00325  -0.0508   0.4325   0.0947
   4.000   0.8413   0.01137   0.00331  -0.0509   0.4282   0.0993
   4.250   0.8684   0.01145   0.00340  -0.0509   0.4243   0.1061
   4.500   0.8959   0.01149   0.00350  -0.0511   0.4206   0.1178
   4.750   0.9147   0.01057   0.00365  -0.0499   0.4166   0.6589
   6.000   1.1371   0.01101   0.00486  -0.0693   0.3925   1.0000
   6.250   1.1617   0.01113   0.00501  -0.0689   0.3884   1.0000
   6.500   1.1858   0.01129   0.00518  -0.0684   0.3844   1.0000
   6.750   1.2095   0.01148   0.00538  -0.0679   0.3801   1.0000
   7.000   1.2329   0.01167   0.00558  -0.0673   0.3749   1.0000
   7.250   1.2561   0.01187   0.00578  -0.0667   0.3677   1.0000
   7.500   1.2783   0.01212   0.00601  -0.0660   0.3602   1.0000
   7.750   1.3009   0.01233   0.00625  -0.0653   0.3527   1.0000
   8.000   1.3222   0.01262   0.00652  -0.0645   0.3458   1.0000
   8.250   1.3442   0.01286   0.00678  -0.0637   0.3379   1.0000
   8.500   1.3641   0.01319   0.00710  -0.0626   0.3280   1.0000
   8.750   1.3835   0.01355   0.00745  -0.0615   0.3177   1.0000
   9.000   1.4027   0.01389   0.00779  -0.0603   0.3074   1.0000
   9.250   1.4196   0.01434   0.00821  -0.0588   0.2945   1.0000
   9.500   1.4332   0.01493   0.00873  -0.0568   0.2765   1.0000
   9.750   1.4440   0.01561   0.00934  -0.0544   0.2582   1.0000
  10.000   1.4488   0.01648   0.01011  -0.0512   0.2346   1.0000
  10.250   1.4390   0.01772   0.01120  -0.0456   0.2069   1.0000
  10.500   1.4256   0.01955   0.01287  -0.0407   0.1766   1.0000
  10.750   1.4110   0.02216   0.01530  -0.0371   0.1431   1.0000
  11.000   1.3841   0.02629   0.01917  -0.0338   0.0959   1.0000
  11.250   1.3577   0.03090   0.02358  -0.0314   0.0565   1.0000
  11.500   1.3395   0.03503   0.02761  -0.0299   0.0309   1.0000
  11.750   1.3354   0.03794   0.03053  -0.0290   0.0233   1.0000
  12.000   1.3356   0.04047   0.03310  -0.0284   0.0204   1.0000
  12.250   1.3353   0.04307   0.03576  -0.0278   0.0185   1.0000
  12.500   1.3368   0.04554   0.03830  -0.0273   0.0174   1.0000
  12.750   1.3381   0.04808   0.04091  -0.0269   0.0165   1.0000
  13.000   1.3385   0.05079   0.04370  -0.0265   0.0158   1.0000
  13.250   1.3381   0.05367   0.04665  -0.0263   0.0152   1.0000
  13.500   1.3362   0.05677   0.04982  -0.0261   0.0145   1.0000
  13.750   1.3352   0.05981   0.05294  -0.0260   0.0140   1.0000
  14.000   1.3358   0.06270   0.05592  -0.0260   0.0136   1.0000
  14.250   1.3356   0.06575   0.05905  -0.0260   0.0132   1.0000
  14.500   1.3345   0.06894   0.06232  -0.0261   0.0128   1.0000
  14.750   1.3332   0.07218   0.06563  -0.0263   0.0124   1.0000
  15.000   1.3313   0.07553   0.06906  -0.0265   0.0120   1.0000
  15.250   1.3280   0.07913   0.07274  -0.0268   0.0117   1.0000
  15.500   1.3240   0.08284   0.07652  -0.0271   0.0114   1.0000
  15.750   1.3187   0.08676   0.08052  -0.0276   0.0112   1.0000
  16.000   1.3140   0.09069   0.08452  -0.0281   0.0110   1.0000
<< Back to NACA M21 AIRFOIL (m21-il)

Polar data table (+)

Polar graphs


<< Back to NACA M21 AIRFOIL (m21-il)