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NACA 0012 AIRFOILS (n0012-il)

NACA 0012 AIRFOILS - NACA 0012 airfoil


Airfoil n0012-il
Details Dat file Parser  
(n0012-il) NACA 0012 AIRFOILS
NACA 0012 airfoil
Max thickness 12% at 30% chord.
Max camber 0% at 0% chord
Source UIUC Airfoil Coordinates Database
Source dat file
The dat file is in Lednicer format
 NACA 0012 AIRFOILS
      66.       66.

 0.0000000 0.0000000
 0.0005839 0.0042603
 0.0023342 0.0084289
 0.0052468 0.0125011
 0.0093149 0.0164706
 0.0145291 0.0203300
 0.0208771 0.0240706
 0.0283441 0.0276827
 0.0369127 0.0311559
 0.0465628 0.0344792
 0.0572720 0.0376414
 0.0690152 0.0406310
 0.0817649 0.0434371
 0.0954915 0.0460489
 0.1101628 0.0484567
 0.1257446 0.0506513
 0.1422005 0.0526251
 0.1594921 0.0543715
 0.1775789 0.0558856
 0.1964187 0.0571640
 0.2159676 0.0582048
 0.2361799 0.0590081
 0.2570083 0.0595755
 0.2784042 0.0599102
 0.3003177 0.0600172
 0.3226976 0.0599028
 0.3454915 0.0595747
 0.3686463 0.0590419
 0.3921079 0.0583145
 0.4158215 0.0574033
 0.4397317 0.0563200
 0.4637826 0.0550769
 0.4879181 0.0536866
 0.5120819 0.0521620
 0.5362174 0.0505161
 0.5602683 0.0487619
 0.5841786 0.0469124
 0.6078921 0.0449802
 0.6313537 0.0429778
 0.6545085 0.0409174
 0.6773025 0.0388109
 0.6996823 0.0366700
 0.7215958 0.0345058
 0.7429917 0.0323294
 0.7638202 0.0301515
 0.7840324 0.0279828
 0.8035813 0.0258337
 0.8224211 0.0237142
 0.8405079 0.0216347
 0.8577995 0.0196051
 0.8742554 0.0176353
 0.8898372 0.0157351
 0.9045085 0.0139143
 0.9182351 0.0121823
 0.9309849 0.0105485
 0.9427280 0.0090217
 0.9534372 0.0076108
 0.9630873 0.0063238
 0.9716559 0.0051685
 0.9791229 0.0041519
 0.9854709 0.0032804
 0.9906850 0.0025595
 0.9947532 0.0019938
 0.9976658 0.0015870
 0.9994161 0.0013419
 1.0000000 0.0012600

 0.0000000 0.0000000
 0.0005839 -.0042603
 0.0023342 -.0084289
 0.0052468 -.0125011
 0.0093149 -.0164706
 0.0145291 -.0203300
 0.0208771 -.0240706
 0.0283441 -.0276827
 0.0369127 -.0311559
 0.0465628 -.0344792
 0.0572720 -.0376414
 0.0690152 -.0406310
 0.0817649 -.0434371
 0.0954915 -.0460489
 0.1101628 -.0484567
 0.1257446 -.0506513
 0.1422005 -.0526251
 0.1594921 -.0543715
 0.1775789 -.0558856
 0.1964187 -.0571640
 0.2159676 -.0582048
 0.2361799 -.0590081
 0.2570083 -.0595755
 0.2784042 -.0599102
 0.3003177 -.0600172
 0.3226976 -.0599028
 0.3454915 -.0595747
 0.3686463 -.0590419
 0.3921079 -.0583145
 0.4158215 -.0574033
 0.4397317 -.0563200
 0.4637826 -.0550769
 0.4879181 -.0536866
 0.5120819 -.0521620
 0.5362174 -.0505161
 0.5602683 -.0487619
 0.5841786 -.0469124
 0.6078921 -.0449802
 0.6313537 -.0429778
 0.6545085 -.0409174
 0.6773025 -.0388109
 0.6996823 -.0366700
 0.7215958 -.0345058
 0.7429917 -.0323294
 0.7638202 -.0301515
 0.7840324 -.0279828
 0.8035813 -.0258337
 0.8224211 -.0237142
 0.8405079 -.0216347
 0.8577995 -.0196051
 0.8742554 -.0176353
 0.8898372 -.0157351
 0.9045085 -.0139143
 0.9182351 -.0121823
 0.9309849 -.0105485
 0.9427280 -.0090217
 0.9534372 -.0076108
 0.9630873 -.0063238
 0.9716559 -.0051685
 0.9791229 -.0041519
 0.9854709 -.0032804
 0.9906850 -.0025595
 0.9947532 -.0019938
 0.9976658 -.0015870
 0.9994161 -.0013419
 1.0000000 -.0012600
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Polars for NACA 0012 AIRFOILS (n0012-il)

PlotAirfoilReynolds #NcritMax Cl/CdDescriptionSource 
   n0012-il50,000925.7 at α=5°Mach=0 Ncrit=9Xfoil predictionDetails
   n0012-il50,000526.5 at α=5.5°Mach=0 Ncrit=5Xfoil predictionDetails
   n0012-il100,000936.7 at α=5°Mach=0 Ncrit=9Xfoil predictionDetails
   n0012-il100,000536.1 at α=5.5°Mach=0 Ncrit=5Xfoil predictionDetails
   n0012-il200,000947.4 at α=5°Mach=0 Ncrit=9Xfoil predictionDetails
   n0012-il200,000545.9 at α=6.5°Mach=0 Ncrit=5Xfoil predictionDetails
   n0012-il500,000961.7 at α=6.5°Mach=0 Ncrit=9Xfoil predictionDetails
   n0012-il500,000561.7 at α=7.5°Mach=0 Ncrit=5Xfoil predictionDetails
   n0012-il1,000,000975.6 at α=7.5°Mach=0 Ncrit=9Xfoil predictionDetails
   n0012-il1,000,000575.4 at α=8.5°Mach=0 Ncrit=5Xfoil predictionDetails
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