SG6040 Airfoil
sg6040Properties
Plot Type
Ncrit
Reynolds Numbers
Kulfan Parameters
Airfoil Geometry
SG6040 at a glance
At a Reynolds number of 1,000,000 the SG6040 reaches a maximum lift coefficient of 1.53 at 18.0° angle of attack, a best lift-to-drag ratio of 114 at Cl 0.99 (5.5°), and a minimum drag coefficient of 0.0062. It carries zero lift at -3.6°.
| Reynolds number | 200,000 | 1,000,000 | 5,000,000 |
|---|---|---|---|
| Max Clthe stall peak | 1.39 | 1.53 | 1.66 |
| at alphaangle of attack it occurs at | 17.5° | 18.0° | 18.0° |
| Best L/Dhighest lift-to-drag ratio | 74 | 114 | 139 |
| at Cllift coefficient it occurs at | 1.15 (8.0°) | 0.99 (5.5°) | 0.86 (4.0°) |
| Min Cdbottom of the drag bucket | 0.0130 | 0.0062 | 0.0046 |
| Zero-lift alphawhere Cl crosses zero in attached flow |
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
No specific aircraft mappings found for "sg6040" in the current database.
Sections like the SG6040
The five closest shapes in the library, by maximum thickness and camber and where along the chord each falls.
- GOE 601 AIRFOIL16.1% thick · 2.5% camber
- HQ 3.0/15 AIRFOIL15.0% thick · 3.0% camber
- NACA 64(2)-41515.0% thick · 2.2% camber
- NACA 64(2)-415 [NACA 6a series from Theory of Wing Sections]15.0% thick · 2.2% camber
- NACA 63-415 AIRFOIL15.0% thick · 2.2% 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.