NACA 65(4)-421 Airfoil
naca654421Properties
Plot Type
Ncrit
Reynolds Numbers
Kulfan Parameters
Airfoil Geometry
NACA 65(4)-421 at a glance
At a Reynolds number of 1,000,000 the NACA 65(4)-421 reaches a best lift-to-drag ratio of 126 at Cl 1.12 (6.5°), and a minimum drag coefficient of 0.0072. It has not stalled by the end of the stored sweep. It carries zero lift at -2.9°.
| Reynolds number | 200,000 | 1,000,000 | 5,000,000 |
|---|---|---|---|
| Max Clthe stall peak | 1.37 | — | 1.70 |
| at alphaangle of attack it occurs at | 18.0° | — | 21.5° |
| Best L/Dhighest lift-to-drag ratio | 54 | 126 | 166 |
| at Cllift coefficient it occurs at | 1.18 (9.0°) | 1.12 (6.5°) | 0.87 (4.0°) |
| Min Cdbottom of the drag bucket | 0.0166 | 0.0072 | 0.0046 |
| Zero-lift alphawhere Cl crosses zero in attached flow | -2.7° |
Read from the stored NeuralFoil sweep at Ncrit 9, sampled every 0.5° of angle of attack. Max Cl is the stall peak rather than the largest value in the sweep, which runs far enough past stall for lift to climb again; a dash means the section had not stalled by the end of it, or for the zero-lift angle, that it stalls negatively before lift reaches zero. Each figure is the sample that won, and nothing is interpolated between Reynolds numbers — the exception is the zero-lift angle, which falls between two samples and is interpolated linearly across the crossing. Load the performance data below for the full curves.
Performance Metrics
These curves are computed with NeuralFoil, a neural network trained on XFOIL, from a stored sweep across Reynolds number, Ncrit and angle of attack. How the polars are computed, and where they stop being trustworthy →
Run it at your Reynolds number
The curves above are a stored sweep at fixed Reynolds numbers. Enter the one your chord and speed actually give and NeuralFoil computes that polar now, onto this chart.
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Aircraft Usage
Found in UIUC Database
Aircraft recorded with two or more sections have a page of their own. All recorded aircraft
Sections like the NACA 65(4)-421
The five closest shapes in the library, by maximum thickness and camber and where along the chord each falls.
- NACA 65(4)-421 a=0.521.0% thick · 3.0% camber
- NACA 65(4)-22121.0% thick · 1.1% camber
- NACA 64(4)-42121.0% thick · 2.2% camber
- WORTMANN FX 05-191 AIRFOIL19.1% thick · 2.6% camber
- NACA 65(3)-61818.0% thick · 3.3% camber
Ranked on geometry alone, from the stored CST parameters — no polar is consulted, so a section that sits close here can still behave differently at the Reynolds number you fly. A near-symmetric section is matched on thickness and its distribution only: with no camber to speak of, where its camber peaks is rounding error rather than shape.