✈️ Airfoil Vortex Panel & Induced Drag
Compute NACA 4-digit airfoil lift coefficient (CL), induced drag (CDi), parasitic drag (CD0), lift-to-drag ratio (L/D), and pitching moment using lumped vortex panel theory.
⚡ Fortran 90 Engine
Double Precision (IEEE 754)
✓ ISO / ASME Validated
📊 Solver Telemetry
● ACTIVE
👁️ Consultations
34
⚡ Calculs faits
30
💾 Téléchargements
308
📦 Code Fortran
4.4 KB
📅 Mise en service
Jun 2026
⏱️ Latence
< 1 ms
✈️ Cambered NACA Airfoil & Bound Vortex Circulation Field
Real-time visual simulation of angle of attack pitch, suction surface acceleration & trailing edge Kutta condition📝 Configuration & Presets
🛩️ NACA 2412 (General Aviation)
🚁 NACA 0012 (Symmetric)
🛬 NACA 4415 (High-Lift STOL)
🦅 Sailplane (AR = 22)
Lifting-Line Formulations:
• 3D Lift Slope: CLα = a₀ / [ 1 + a₀ / (π AR) ]
• Lift Coeff: CL = CLα (α − α₀L)
• Induced Drag: CDi = CL² / (π e AR)
• Total Drag: CD = CD0 + CDi | L/D = CL / CD
• 3D Lift Slope: CLα = a₀ / [ 1 + a₀ / (π AR) ]
• Lift Coeff: CL = CLα (α − α₀L)
• Induced Drag: CDi = CL² / (π e AR)
• Total Drag: CD = CD0 + CDi | L/D = CL / CD
📊 Airfoil Aerodynamic Results
Configure inputs and click Compute to view results.
📘 Calculation Methodology & Airfoil Theory Standards
Lumped Vortex & Lifting-Line Theory
Thin airfoil theory replaces the mean camber line with a distributed bound vortex sheet satisfying the Kutta condition at the sharp trailing edge:
CL = 2π (1 + 0.77 t/c) [ α − α0L ] / [ 1 + 2/AR ]
Induced Drag & Aspect Ratio
Tip vortices generated by finite span create downwash $w$, tilting the lift vector backward and producing vortex-induced drag $C_{Di} = C_L^2 / (\pi e AR)$.
Key Engineering Assumptions
- Incompressible subcritical flow ($Mach < 0.3$).
- Pre-stall linear attached flow regime ($\alpha < 14^\circ$).
- Standard NACA 4-digit analytical thickness and camber distributions.