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NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA M25 AIRFOIL (m25-il)
Reynolds number: 500,000
Max Cl/Cd: 102.28 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m25-il-500000.txt
Download as CSV file: xf-m25-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M25 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.2940   0.12857   0.12568   0.0204   0.6413   0.0092
 -10.250  -0.2892   0.12553   0.12260   0.0195   0.6376   0.0097
  -7.750  -0.3396   0.10804   0.10495   0.0200   0.6132   0.0106
  -7.500  -0.3331   0.10398   0.10084   0.0196   0.6095   0.0107
  -7.250  -0.3250   0.10089   0.09773   0.0180   0.6057   0.0109
  -7.000  -0.3136   0.09780   0.09464   0.0156   0.6015   0.0110
  -6.750  -0.3005   0.09474   0.09156   0.0130   0.5976   0.0113
  -6.500  -0.2861   0.09168   0.08844   0.0102   0.5942   0.0116
  -6.250  -0.2701   0.08855   0.08528   0.0073   0.5906   0.0119
  -6.000  -0.2527   0.08536   0.08208   0.0043   0.5867   0.0124
  -5.750  -0.2338   0.08213   0.07881   0.0011   0.5830   0.0130
  -5.500  -0.2107   0.07902   0.07562  -0.0026   0.5795   0.0136
  -5.250  -0.1814   0.07623   0.07277  -0.0076   0.5760   0.0138
  -5.000  -0.1503   0.07343   0.06988  -0.0128   0.5724   0.0139
  -4.500  -0.1169   0.06537   0.06173  -0.0156   0.5656   0.0143
  -4.250  -0.0952   0.06265   0.05895  -0.0175   0.5621   0.0150
  -4.000  -0.0696   0.05993   0.05618  -0.0200   0.5584   0.0159
  -3.750  -0.0418   0.05711   0.05328  -0.0226   0.5550   0.0168
  -3.500  -0.0061   0.05490   0.05093  -0.0258   0.5516   0.0177
  -3.250   0.0322   0.05290   0.04875  -0.0293   0.5483   0.0179
  -3.000   0.0670   0.05039   0.04611  -0.0318   0.5447   0.0180
  -2.750   0.0858   0.04591   0.04159  -0.0326   0.5416   0.0183
  -2.500   0.1057   0.04327   0.03890  -0.0333   0.5383   0.0189
  -2.250   0.1323   0.04130   0.03681  -0.0344   0.5349   0.0199
  -2.000   0.1621   0.03920   0.03464  -0.0357   0.5316   0.0212
  -1.750   0.2044   0.03843   0.03365  -0.0372   0.5280   0.0230
  -1.500   0.2368   0.03662   0.03168  -0.0381   0.5248   0.0232
  -1.250   0.2674   0.03477   0.02964  -0.0387   0.5217   0.0233
  -1.000   0.2881   0.03093   0.02573  -0.0393   0.5186   0.0241
  -0.750   0.3138   0.02927   0.02402  -0.0399   0.5151   0.0249
  -0.500   0.3420   0.02785   0.02249  -0.0403   0.5118   0.0262
   0.500   0.4649   0.02219   0.01617  -0.0411   0.4989   0.0316
   0.750   0.4913   0.02109   0.01500  -0.0414   0.4956   0.0327
   1.000   0.5190   0.02024   0.01401  -0.0417   0.4925   0.0342
   1.250   0.5480   0.01941   0.01310  -0.0418   0.4894   0.0368
   1.500   0.5797   0.01923   0.01277  -0.0416   0.4858   0.0403
   1.750   0.6080   0.01731   0.01065  -0.0416   0.4825   0.0420
   2.000   0.6351   0.01667   0.00995  -0.0419   0.4792   0.0438
   2.250   0.6632   0.01617   0.00937  -0.0421   0.4760   0.0474
   2.500   0.6925   0.01541   0.00847  -0.0421   0.4727   0.0550
   2.750   0.7202   0.01503   0.00810  -0.0424   0.4692   0.0602
   3.000   0.7486   0.01454   0.00749  -0.0425   0.4660   0.0690
   3.750   0.8345   0.01268   0.00530  -0.0423   0.4557   0.0552
   4.000   0.8617   0.01294   0.00569  -0.0430   0.4522   0.0818
   4.250   0.8888   0.01187   0.00435  -0.0420   0.4493   0.0423
   4.500   0.9161   0.01172   0.00424  -0.0420   0.4460   0.0418
   4.750   0.9431   0.01155   0.00410  -0.0420   0.4424   0.0430
   5.000   0.9705   0.01149   0.00405  -0.0422   0.4381   0.0453
   5.250   0.9980   0.01147   0.00406  -0.0424   0.4314   0.0480
   5.500   1.0258   0.01144   0.00405  -0.0426   0.4238   0.0519
   5.750   1.0535   0.01141   0.00405  -0.0429   0.4126   0.0592
   6.000   1.0807   0.01138   0.00419  -0.0432   0.4015   0.1809
   6.500   1.1752   0.01149   0.00509  -0.0540   0.2972   1.0000
   6.750   1.1826   0.01644   0.00862  -0.0555   0.0277   1.0000
   7.000   1.2048   0.01702   0.00928  -0.0551   0.0236   1.0000
   7.250   1.2255   0.01776   0.01011  -0.0547   0.0209   1.0000
   7.500   1.2433   0.01882   0.01133  -0.0540   0.0182   1.0000
   7.750   1.2579   0.02009   0.01274  -0.0532   0.0173   1.0000
   8.000   1.2721   0.02123   0.01397  -0.0524   0.0168   1.0000
   8.250   1.2816   0.02280   0.01565  -0.0516   0.0163   1.0000
   8.500   1.2818   0.02510   0.01806  -0.0509   0.0159   1.0000
   8.750   1.2764   0.02760   0.02067  -0.0492   0.0157   1.0000
   9.000   1.2737   0.03036   0.02352  -0.0484   0.0154   1.0000
   9.250   1.2717   0.03321   0.02646  -0.0479   0.0149   1.0000
   9.500   1.2693   0.03614   0.02949  -0.0474   0.0143   1.0000
   9.750   1.2664   0.03914   0.03255  -0.0469   0.0138   1.0000
  10.000   1.2622   0.04228   0.03577  -0.0463   0.0136   1.0000
  10.250   1.2578   0.04541   0.03897  -0.0457   0.0134   1.0000
  10.500   1.2541   0.04842   0.04205  -0.0450   0.0132   1.0000
  10.750   1.2514   0.05124   0.04493  -0.0442   0.0130   1.0000
  11.000   1.2507   0.05384   0.04757  -0.0433   0.0128   1.0000
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