NACA 0012 AIRFOILS (n0012-il)
NACA 0012 AIRFOILS - NACA 0012 airfoil
| Details | Dat file | Parser | |
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(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)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n0012-il | 50,000 | 9 | 25.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 50,000 | 5 | 26.5 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 100,000 | 9 | 36.7 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 100,000 | 5 | 36.1 at α=5.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 200,000 | 9 | 47.4 at α=5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 200,000 | 5 | 45.9 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 500,000 | 9 | 61.7 at α=6.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 500,000 | 5 | 61.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n0012-il | 1,000,000 | 9 | 75.6 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n0012-il | 1,000,000 | 5 | 75.4 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
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