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

NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 50,000
Max Cl/Cd: 20.63 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m1-il-50000-n5.txt
Download as CSV file: xf-m1-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-M1 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5665   0.11358   0.10688   0.0100   1.0000   0.0584
 -10.000  -0.5684   0.10914   0.10248   0.0088   1.0000   0.0574
  -9.750  -0.7074   0.11293   0.10608   0.0153   1.0000   0.0482
  -9.500  -0.7026   0.10855   0.10171   0.0154   1.0000   0.0472
  -9.250  -0.7022   0.10370   0.09690   0.0133   1.0000   0.0466
  -9.000  -0.7032   0.09859   0.09184   0.0103   1.0000   0.0463
  -8.750  -0.7053   0.09318   0.08647   0.0065   1.0000   0.0459
  -8.500  -0.7091   0.08768   0.08101   0.0024   1.0000   0.0454
  -8.250  -0.7118   0.08220   0.07549  -0.0010   1.0000   0.0451
  -8.000  -0.7132   0.07675   0.06996  -0.0039   1.0000   0.0449
  -7.750  -0.7130   0.07138   0.06444  -0.0062   1.0000   0.0450
  -7.500  -0.7103   0.06609   0.05892  -0.0080   1.0000   0.0452
  -7.250  -0.7047   0.06091   0.05336  -0.0092   1.0000   0.0454
  -7.000  -0.6960   0.05588   0.04793  -0.0098   1.0000   0.0455
  -6.750  -0.6841   0.05106   0.04263  -0.0100   1.0000   0.0456
  -6.500  -0.6690   0.04654   0.03756  -0.0098   1.0000   0.0461
  -6.250  -0.6508   0.04261   0.03280  -0.0093   1.0000   0.0484
  -6.000  -0.6323   0.03927   0.02932  -0.0091   1.0000   0.0521
  -5.750  -0.6102   0.03613   0.02572  -0.0085   1.0000   0.0547
  -5.500  -0.5860   0.03327   0.02228  -0.0078   1.0000   0.0586
  -5.250  -0.5616   0.03084   0.01941  -0.0072   1.0000   0.0650
  -5.000  -0.5359   0.02865   0.01689  -0.0065   1.0000   0.0693
  -4.750  -0.5094   0.02696   0.01479  -0.0058   1.0000   0.0768
  -4.500  -0.4842   0.02536   0.01315  -0.0052   1.0000   0.0845
  -4.250  -0.4577   0.02395   0.01147  -0.0043   1.0000   0.0886
  -4.000  -0.4324   0.02265   0.01006  -0.0036   1.0000   0.0937
  -3.750  -0.4067   0.02156   0.00875  -0.0030   1.0000   0.1024
  -3.500  -0.3819   0.02043   0.00762  -0.0025   1.0000   0.1187
  -3.250  -0.3586   0.01896   0.00654  -0.0020   1.0000   0.1788
  -3.000  -0.3437   0.01666   0.00593  -0.0001   1.0000   0.4845
  -2.750  -0.3203   0.01563   0.00623   0.0047   1.0000   0.8387
  -2.500  -0.2117   0.01581   0.00581  -0.0090   1.0000   1.0000
  -2.250  -0.1909   0.01552   0.00526  -0.0084   1.0000   1.0000
  -2.000  -0.1700   0.01529   0.00482  -0.0076   1.0000   1.0000
  -1.750  -0.1489   0.01510   0.00445  -0.0068   1.0000   1.0000
  -1.500  -0.1277   0.01495   0.00414  -0.0059   1.0000   1.0000
  -1.250  -0.1064   0.01483   0.00387  -0.0050   1.0000   1.0000
  -1.000  -0.0851   0.01473   0.00367  -0.0041   1.0000   1.0000
  -0.750  -0.0639   0.01466   0.00352  -0.0031   1.0000   1.0000
  -0.500  -0.0426   0.01461   0.00341  -0.0021   1.0000   1.0000
  -0.250  -0.0213   0.01458   0.00334  -0.0010   1.0000   1.0000
   0.000   0.0000   0.01457   0.00331   0.0000   1.0000   1.0000
   0.250   0.0213   0.01458   0.00334   0.0010   1.0000   1.0000
   0.500   0.0426   0.01461   0.00341   0.0021   1.0000   1.0000
   0.750   0.0639   0.01466   0.00351   0.0031   1.0000   1.0000
   1.000   0.0851   0.01473   0.00367   0.0041   1.0000   1.0000
   1.250   0.1064   0.01482   0.00387   0.0050   1.0000   1.0000
   1.500   0.1277   0.01495   0.00414   0.0059   1.0000   1.0000
   1.750   0.1489   0.01510   0.00445   0.0068   1.0000   1.0000
   2.000   0.1700   0.01529   0.00482   0.0076   1.0000   1.0000
   2.250   0.1910   0.01552   0.00526   0.0084   1.0000   1.0000
   2.500   0.2118   0.01580   0.00581   0.0090   1.0000   1.0000
   2.750   0.3204   0.01563   0.00623  -0.0047   0.8381   1.0000
   3.000   0.3437   0.01666   0.00593   0.0001   0.4837   1.0000
   3.250   0.3586   0.01897   0.00654   0.0020   0.1783   1.0000
   3.500   0.3819   0.02043   0.00762   0.0025   0.1187   1.0000
   3.750   0.4068   0.02156   0.00875   0.0030   0.1024   1.0000
   4.000   0.4324   0.02265   0.01006   0.0036   0.0937   1.0000
   4.250   0.4578   0.02395   0.01147   0.0043   0.0886   1.0000
   4.500   0.4842   0.02536   0.01315   0.0052   0.0845   1.0000
   4.750   0.5094   0.02695   0.01479   0.0058   0.0768   1.0000
   5.000   0.5359   0.02865   0.01690   0.0065   0.0693   1.0000
   5.250   0.5617   0.03084   0.01941   0.0072   0.0650   1.0000
   5.500   0.5860   0.03327   0.02228   0.0078   0.0586   1.0000
   5.750   0.6102   0.03613   0.02571   0.0085   0.0547   1.0000
   6.000   0.6323   0.03927   0.02932   0.0091   0.0521   1.0000
   6.250   0.6508   0.04261   0.03280   0.0093   0.0483   1.0000
   6.500   0.6691   0.04654   0.03756   0.0098   0.0461   1.0000
   6.750   0.6841   0.05106   0.04264   0.0100   0.0456   1.0000
   7.000   0.6960   0.05589   0.04794   0.0098   0.0455   1.0000
   7.250   0.7048   0.06092   0.05336   0.0092   0.0454   1.0000
   7.500   0.7104   0.06610   0.05893   0.0079   0.0451   1.0000
   7.750   0.7130   0.07139   0.06445   0.0062   0.0450   1.0000
   8.000   0.7134   0.07677   0.06998   0.0038   0.0449   1.0000
   8.250   0.7120   0.08221   0.07551   0.0009   0.0451   1.0000
   8.500   0.7094   0.08771   0.08103  -0.0025   0.0454   1.0000
   8.750   0.7056   0.09322   0.08651  -0.0066   0.0459   1.0000
   9.000   0.7036   0.09863   0.09188  -0.0104   0.0463   1.0000
   9.250   0.7027   0.10373   0.09693  -0.0134   0.0466   1.0000
   9.500   0.7031   0.10858   0.10175  -0.0155   0.0472   1.0000
   9.750   0.7081   0.11295   0.10611  -0.0153   0.0482   1.0000
  10.000   0.5678   0.10895   0.10229  -0.0087   0.0575   1.0000
  10.250   0.5660   0.11337   0.10668  -0.0099   0.0585   1.0000
<< Back to NACA-M1 AIRFOIL (m1-il)

Polar data table (+)

Polar graphs


<< Back to NACA-M1 AIRFOIL (m1-il)