NACA 65(4)-421 a=0.5 (naca654421a05-il)
NACA 65(4)-421 a=0.5 - NACA 65(4)-421 a=0.5 airfoil
| Details | Dat file | Parser | |
|
(naca654421a05-il) NACA 65(4)-421 a=0.5 NACA 65(4)-421 a=0.5 airfoil Max thickness 21% at 39.9% chord. Max camber 3% at 45% chord Source UIUC Airfoil Coordinates Database Source dat file The dat file is in Selig format |
NACA 65(4)-421 a=0.5 1.00000 0.00000 0.95034 0.00833 0.90100 0.02095 0.85181 0.03550 0.80264 0.05097 0.75340 0.06664 0.70398 0.08193 0.65428 0.09637 0.60421 0.10942 0.55362 0.12056 0.50211 0.12890 0.45026 0.13362 0.39882 0.13470 0.34754 0.13258 0.29639 0.12765 0.24538 0.12000 0.19455 0.10951 0.14395 0.09572 0.09371 0.07773 0.06880 0.06651 0.04414 0.05315 0.01992 0.03631 0.00813 0.02553 0.00363 0.01991 0.00155 0.01620 0.00000 0.00000 0.00845 -0.01344 0.01137 -0.01603 0.01687 -0.01965 0.03008 -0.02595 0.05586 -0.03551 0.08120 -0.04275 0.10629 -0.04869 0.15605 -0.05780 0.20545 -0.06455 0.25462 -0.06952 0.30361 -0.07293 0.35246 -0.07486 0.40118 -0.07526 0.44974 -0.07370 0.49789 -0.07010 0.54638 -0.06484 0.59579 -0.05818 0.64572 -0.05057 0.69602 -0.04229 0.74660 -0.03360 0.79736 -0.02485 0.84819 -0.01634 0.89900 -0.00867 0.94966 -0.00257 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 65(4)-421 a=0.5 (naca654421a05-il)
| Plot | Airfoil | Reynolds # | Ncrit | Max Cl/Cd | Description | Source | |
|---|---|---|---|---|---|---|---|
| naca654421a05-il | 50,000 | 9 | 3.7 at α=12.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 50,000 | 5 | 10.2 at α=14.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 100,000 | 9 | 17.4 at α=13.5° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 100,000 | 5 | 16.5 at α=12.75° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 200,000 | 9 | 31.7 at α=11.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 200,000 | 5 | 44.4 at α=6.5° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 500,000 | 9 | 96.5 at α=7.75° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 500,000 | 5 | 92.5 at α=7.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| naca654421a05-il | 1,000,000 | 9 | 128.2 at α=7.25° | Mach=0 Ncrit=9 | Xfoil prediction | Details | |
| naca654421a05-il | 1,000,000 | 5 | 117.6 at α=6.25° | Mach=0 Ncrit=5 | Xfoil prediction | Details | |
| Reynolds number calculator | |||||||