NACA 0012 AIRFOILS (n0012

NACA 0012 AIRFOILS - NACA 0012 airfoil

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 No parser warnings Send to airfoil plotter
Add to comparison
Lednicer format dat file
Selig format dat file

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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 Reynolds number calculator Set Reynolds number and Ncrit rangeLowHigh Reynolds Number NCrit

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