luokai0 Aerodynamics Expert
Before Starting
- Subsonic, transonic, or supersonic regime?
- Internal or external flow?
- Analysis or design focus?
Core Expertise Areas
Airfoil Theory
Lift: generated by pressure difference between upper and lower surfaces. Camber: curvature of mean line, increases lift at zero angle of attack. Thickness: affects drag and maximum lift, NACA 4-digit series defines profile. Angle of attack: increasing AoA increases lift until stall. Stall: boundary layer separates from suction surface, lift drops suddenly.
Thin Airfoil Theory
Lift coefficient: CL = 2 pi times alpha for thin symmetric airfoil. Moment coefficient: CM about quarter chord is zero for symmetric airfoils. Aerodynamic center: point where moment coefficient is independent of AoA. Camber effect: adds lift at zero AoA proportional to maximum camber.
Drag
Pressure drag: form drag from pressure distribution, reduced by streamlining. Skin friction drag: viscous shear stress on surface, dominant for streamlined bodies. Induced drag: due to finite wing span, CDi = CL squared over pi AR e. Wave drag: energy lost to shock waves in transonic and supersonic flow. Drag polar: CD vs CL squared, slope is 1 over pi AR e.
Compressible Flow
Mach number: M = V over a, ratio of flow speed to speed of sound. Critical Mach: freestream Mach where local sonic flow first appears. Prandtl-Glauert: compressibility correction for subsonic flow, 1 over sqrt 1 minus M squared. Shock waves: normal and oblique, pressure rises discontinuously. Expansion fans: isentropic acceleration around convex corners.
Best Practices
- Verify Reynolds number and Mach number before selecting analysis method
- Use panel methods for subsonic, Euler for transonic, full NS for separated flow
- Validate CFD results against wind tunnel data for critical designs
- Check for flow separation before assuming attached flow methods are valid
Common Pitfalls
| Pitfall | Fix |
|---|---|
| Applying thin airfoil theory at high AoA | Valid only for small angles, below stall |
| Ignoring compressibility near Mach 0.3 | Apply Prandtl-Glauert correction above M=0.3 |
| 2D analysis for finite wing | Apply finite wing correction for induced drag |
| Inviscid analysis near separation | Use viscous solver for high AoA or bluff bodies |
Related Skills
- flight-mechanics-expert
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