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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 100,000
Max Cl/Cd: 45.26 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m10-il-100000-n5.txt
Download as CSV file: xf-m10-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M10 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  -0.4576   0.11385   0.10916   0.0031   1.0000   0.0406
  -9.750  -0.4611   0.11048   0.10582   0.0009   1.0000   0.0412
  -9.500  -0.5800   0.11726   0.11234   0.0160   1.0000   0.0332
  -9.250  -0.5750   0.11373   0.10883   0.0150   1.0000   0.0348
  -9.000  -0.5715   0.10999   0.10511   0.0133   1.0000   0.0357
  -8.750  -0.5686   0.10605   0.10120   0.0114   1.0000   0.0357
  -8.500  -0.5666   0.10184   0.09703   0.0090   1.0000   0.0347
  -8.250  -0.5672   0.09670   0.09193   0.0040   1.0000   0.0293
  -8.000  -0.5642   0.09259   0.08785   0.0028   1.0000   0.0283
  -7.750  -0.5620   0.08829   0.08358  -0.0001   1.0000   0.0276
  -7.500  -0.5566   0.08356   0.07886  -0.0040   1.0000   0.0270
  -7.250  -0.5494   0.07862   0.07392  -0.0081   1.0000   0.0265
  -7.000  -0.5401   0.07352   0.06878  -0.0122   1.0000   0.0261
  -6.750  -0.5288   0.06829   0.06346  -0.0161   1.0000   0.0259
  -6.500  -0.5149   0.06302   0.05807  -0.0196   1.0000   0.0265
  -6.250  -0.4985   0.05757   0.05243  -0.0229   1.0000   0.0277
  -6.000  -0.4801   0.05188   0.04646  -0.0254   1.0000   0.0287
  -5.750  -0.4610   0.04622   0.04045  -0.0270   1.0000   0.0290
  -5.500  -0.4405   0.04067   0.03443  -0.0279   1.0000   0.0297
  -5.250  -0.4200   0.03576   0.02900  -0.0282   1.0000   0.0315
  -5.000  -0.3988   0.03458   0.02773  -0.0283   1.0000   0.0352
  -4.750  -0.3751   0.03086   0.02348  -0.0281   1.0000   0.0374
  -4.500  -0.3493   0.02770   0.01954  -0.0275   1.0000   0.0415
  -4.250  -0.3257   0.02534   0.01688  -0.0273   1.0000   0.0448
  -4.000  -0.3007   0.02373   0.01500  -0.0269   1.0000   0.0478
  -3.750  -0.2750   0.02244   0.01330  -0.0264   1.0000   0.0528
  -3.500  -0.2491   0.02087   0.01136  -0.0258   1.0000   0.0541
  -3.250  -0.2237   0.01948   0.00973  -0.0251   1.0000   0.0556
  -3.000  -0.1992   0.01827   0.00847  -0.0246   1.0000   0.0579
  -2.750  -0.1749   0.01743   0.00758  -0.0240   1.0000   0.0603
  -2.500  -0.1508   0.01685   0.00691  -0.0233   1.0000   0.0649
  -2.250  -0.1270   0.01638   0.00633  -0.0226   1.0000   0.0690
  -2.000  -0.1039   0.01565   0.00564  -0.0218   1.0000   0.0708
  -1.750  -0.0811   0.01512   0.00511  -0.0211   1.0000   0.0732
  -1.500  -0.0583   0.01473   0.00472  -0.0203   1.0000   0.0768
  -1.250  -0.0218   0.01429   0.00424  -0.0224   0.9908   0.0848
  -1.000   0.0172   0.01375   0.00382  -0.0249   0.9795   0.1128
  -0.750   0.0524   0.01234   0.00355  -0.0273   0.9679   0.4100
  -0.500   0.1029   0.01089   0.00357  -0.0307   0.9622   1.0000
  -0.250   0.1406   0.01091   0.00342  -0.0329   0.9445   1.0000
   0.000   0.1772   0.01093   0.00331  -0.0347   0.9264   1.0000
   0.250   0.2106   0.01096   0.00324  -0.0357   0.9064   1.0000
   0.500   0.2415   0.01100   0.00319  -0.0361   0.8859   1.0000
   0.750   0.2698   0.01107   0.00318  -0.0359   0.8651   1.0000
   1.000   0.2966   0.01115   0.00319  -0.0354   0.8444   1.0000
   1.250   0.3225   0.01125   0.00324  -0.0347   0.8241   1.0000
   1.500   0.3480   0.01137   0.00331  -0.0339   0.8040   1.0000
   1.750   0.3735   0.01149   0.00339  -0.0331   0.7851   1.0000
   2.000   0.3991   0.01163   0.00354  -0.0324   0.7652   1.0000
   2.250   0.4248   0.01178   0.00368  -0.0317   0.7461   1.0000
   2.500   0.4504   0.01194   0.00384  -0.0310   0.7273   1.0000
   2.750   0.4761   0.01212   0.00404  -0.0303   0.7072   1.0000
   3.000   0.5017   0.01230   0.00427  -0.0296   0.6873   1.0000
   3.250   0.5268   0.01248   0.00448  -0.0287   0.6612   1.0000
   3.500   0.5509   0.01268   0.00462  -0.0276   0.6236   1.0000
   3.750   0.5743   0.01293   0.00477  -0.0262   0.5732   1.0000
   4.000   0.5977   0.01326   0.00493  -0.0251   0.5108   1.0000
   4.250   0.6210   0.01372   0.00518  -0.0240   0.4369   1.0000
   4.500   0.6435   0.01444   0.00556  -0.0231   0.3403   1.0000
   4.750   0.6611   0.01638   0.00637  -0.0222   0.1325   1.0000
   5.000   0.6825   0.01786   0.00751  -0.0216   0.0761   1.0000
   5.250   0.7056   0.01893   0.00862  -0.0209   0.0598   1.0000
   5.500   0.7277   0.02013   0.00992  -0.0201   0.0494   1.0000
   5.750   0.7489   0.02144   0.01127  -0.0193   0.0406   1.0000
   6.000   0.7706   0.02282   0.01281  -0.0182   0.0360   1.0000
   6.250   0.7919   0.02429   0.01433  -0.0174   0.0310   1.0000
   6.500   0.8142   0.02613   0.01631  -0.0164   0.0282   1.0000
   6.750   0.8378   0.02819   0.01864  -0.0154   0.0262   1.0000
   7.000   0.8607   0.03016   0.02090  -0.0145   0.0234   1.0000
   7.250   0.8821   0.03211   0.02311  -0.0138   0.0211   1.0000
   7.500   0.9019   0.03478   0.02609  -0.0128   0.0201   1.0000
   7.750   0.9188   0.03814   0.02982  -0.0118   0.0194   1.0000
   8.000   0.9318   0.04230   0.03444  -0.0106   0.0190   1.0000
   8.250   0.9406   0.04671   0.03941  -0.0093   0.0186   1.0000
   8.500   0.9495   0.04994   0.04317  -0.0079   0.0180   1.0000
   8.750   0.9535   0.05377   0.04750  -0.0066   0.0174   1.0000
   9.000   0.9524   0.05803   0.05217  -0.0054   0.0171   1.0000
   9.250   0.9460   0.06245   0.05692  -0.0047   0.0169   1.0000
   9.500   0.9343   0.06680   0.06153  -0.0042   0.0169   1.0000
   9.750   0.9179   0.07115   0.06605  -0.0044   0.0170   1.0000
  10.000   0.9008   0.07644   0.07148  -0.0071   0.0171   1.0000
  10.250   0.8845   0.08291   0.07805  -0.0118   0.0173   1.0000
  10.500   0.8698   0.09046   0.08567  -0.0176   0.0176   1.0000
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