NACA 63(4)-421 (naca634421-il)
NACA 63(4)-421 - NACA 63(4)-421 airfoil
| Details | Dat file | Parser | |
|
(naca634421-il) NACA 63(4)-421 NACA 63(4)-421 airfoil Max thickness 21% at 34.8% chord. Max camber 2.2% at 50% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
NACA 63(4)-421 1.00000 0.00000 0.95037 0.01022 0.90078 0.02144 0.85111 0.03364 0.80135 0.04643 0.75145 0.05947 0.70143 0.07232 0.65126 0.08455 0.60096 0.09582 0.55054 0.10580 0.50000 0.11412 0.44937 0.12044 0.39567 0.12439 0.34793 0.12558 0.29719 0.12352 0.24649 0.11837 0.19585 0.10993 0.14535 0.09774 0.09506 0.08097 0.07007 0.07010 0.04527 0.05675 0.02086 0.03925 0.00902 0.02717 0.00452 0.02054 0.00237 0.01661 0.00000 0.00000 0.00763 -0.01461 0.01048 -0.01774 0.01598 -0.02289 0.02914 -0.03181 0.05473 -0.04411 0.07993 -0.05314 0.10494 -0.06029 0.15465 -0.07082 0.20415 -0.07809 0.25351 -0.08257 0.30281 -0.08464 0.35207 -0.08438 0.40133 -0.08155 0.45063 -0.07664 0.50000 -0.07000 0.54946 -0.06200 0.59904 -0.05298 0.64874 -0.04335 0.69857 -0.03344 0.74855 -0.02367 0.79865 -0.01459 0.84889 -0.00672 0.89922 -0.00076 0.94963 0.00242 1.00000 0.00000 |
No parser warnings |
Send to airfoil plotter Add to comparison Lednicer format dat file Selig format dat file |
Similar airfoils
|
Polars for NACA 63(4)-421 (naca634421-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca634421-il | 50,000 | 9 | 3.8 at α=12.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634421-il | 50,000 | 5 | 9.4 at α=14.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634421-il | 100,000 | 9 | 24.4 at α=13.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634421-il | 100,000 | 5 | 35 at α=8.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634421-il | 200,000 | 9 | 61.4 at α=9.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634421-il | 200,000 | 5 | 63 at α=8° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634421-il | 500,000 | 9 | 95.3 at α=8° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634421-il | 500,000 | 5 | 90.8 at α=7.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca634421-il | 1,000,000 | 9 | 119.1 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca634421-il | 1,000,000 | 5 | 109.7 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||