NACA 8-H-12 AIRFOIL (n8h12-il)
NACA 8-H-12 AIRFOIL - NACA 8-H-12 rotorcraft airfoil
| Details | Dat file | Parser | |
|
(n8h12-il) NACA 8-H-12 AIRFOIL NACA 8-H-12 rotorcraft airfoil Max thickness 12% at 30% chord. Max camber 3.5% at 30% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Lednicer format |
NACA 8-H-12 AIRFOIL
19. 18.
0.0000000 0.0000000
0.0036000 0.0152000
0.0080000 0.0201000
0.0198000 0.0294000
0.0442000 0.0431000
0.0691000 0.0538000
0.0943000 0.0626000
0.1450000 0.0763000
0.1961000 0.0860000
0.2475000 0.0924000
0.2997000 0.0953000
0.4029000 0.0903000
0.5039000 0.0767000
0.6036000 0.0585000
0.7025000 0.0384000
0.8011000 0.0189000
0.9001999 0.0034000
0.9499000 0.0012000
1.0000000 0.0000000
0.0000000 0.0000000
0.0114000 -.0095000
0.0169000 -.0112000
0.0302000 -.0141000
0.0558000 -.0174000
0.0809000 -.0192000
0.1057000 -.0206000
0.1550000 -.0224000
0.2039000 -.0235000
0.2525000 -.0242000
0.3003000 -.0245000
0.3971000 -.0249000
0.4961000 -.0244000
0.5964000 -.0229000
0.7988000 -.0164000
0.8998001 -.0105000
0.9500000 -.0063000
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 8-H-12 AIRFOIL (n8h12-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| n8h12-il | 50,000 | 9 | 8 at α=-0.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n8h12-il | 50,000 | 5 | 18.6 at α=3.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n8h12-il | 100,000 | 9 | 45.6 at α=9.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n8h12-il | 100,000 | 5 | 48 at α=8.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n8h12-il | 200,000 | 9 | 72 at α=8.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n8h12-il | 200,000 | 5 | 71.2 at α=7.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n8h12-il | 500,000 | 9 | 104.2 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n8h12-il | 500,000 | 5 | 98.6 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| n8h12-il | 1,000,000 | 9 | 128.4 at α=7° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| n8h12-il | 1,000,000 | 5 | 118.2 at α=5.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||