NACA M18 AIRFOIL (m18-il)
NACA M18 AIRFOIL - NACA/Munk M-18 airfoil
| Details | Dat file | Parser | |
|
(m18-il) NACA M18 AIRFOIL NACA/Munk M-18 airfoil Max thickness 12% at 30% chord. Max camber 4.1% at 30% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NACA M18 AIRFOIL
17. 17.
0.0000000 0.0000000
0.0125000 0.0204600
0.0250000 0.0304200
0.0500000 0.0456400
0.0750000 0.0576700
0.1000000 0.0670900
0.1500000 0.0822400
0.2000000 0.0921800
0.3000000 0.1018700
0.4000000 0.0993600
0.5000000 0.0897500
0.6000000 0.0743400
0.7000000 0.0567300
0.8000000 0.0378200
0.9000000 0.0197100
0.9500000 0.0116600
1.0000000 0.0044000
0.0000000 0.0000000
0.0125000 -.0164400
0.0250000 -.0195800
0.0500000 -.0218500
0.0750000 -.0222300
0.1000000 -.0220100
0.1500000 -.0207600
0.2000000 -.0195200
0.3000000 -.0183300
0.4000000 -.0191400
0.5000000 -.0216500
0.6000000 -.0235600
0.7000000 -.0236700
0.8000000 -.0204800
0.9000000 -.0130900
0.9500000 -.0074400
1.0000000 0.0000000
|
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for NACA M18 AIRFOIL (m18-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| m18-il | 50,000 | 9 | 11.1 at α=0.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m18-il | 50,000 | 5 | 25.3 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m18-il | 100,000 | 9 | 47.3 at α=10.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m18-il | 100,000 | 5 | 50.4 at α=9° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m18-il | 200,000 | 9 | 71.8 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m18-il | 200,000 | 5 | 72.2 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m18-il | 500,000 | 9 | 103.5 at α=8.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m18-il | 500,000 | 5 | 100.7 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| m18-il | 1,000,000 | 9 | 127.2 at α=7.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| m18-il | 1,000,000 | 5 | 116.8 at α=7° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||