โ๏ธ 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.
โก Fortran 90 Engine
Double Precision (IEEE 754)
โ ISO / ASME Validated
Cfd
๐ Solver Telemetry
โ ACTIVE
๐๏ธ Views
34
โก Solves
30
๐พ Downloads
407
๐ฆ Fortran Code
4.4 KB
๐
Released
Jun 2026
โฑ๏ธ Latency
< 1 ms
โ๏ธ Cambered NACA Airfoil & Bound Vortex Circulation Field
Real-time visual simulation of angle of attack pitch, suction surface acceleration & trailing edge Kutta condition๐ Configuration & Presets
๐ฉ๏ธ NACA 2412 (General Aviation)
๐ NACA 0012 (Symmetric)
๐ฌ NACA 4415 (High-Lift STOL)
๐ฆ
Sailplane (AR = 22)
Lifting-Line Formulations:
โข 3D Lift Slope: CLฮฑ = aโ / [ 1 + aโ / (ฯ AR) ]
โข Lift Coeff: CL = CLฮฑ (ฮฑ โ ฮฑโL)
โข Induced Drag: CDi = CLยฒ / (ฯ e AR)
โข Total Drag: CD = CD0 + CDi | L/D = CL / CD
โข 3D Lift Slope: CLฮฑ = aโ / [ 1 + aโ / (ฯ AR) ]
โข Lift Coeff: CL = CLฮฑ (ฮฑ โ ฮฑโL)
โข Induced Drag: CDi = CLยฒ / (ฯ e AR)
โข Total Drag: CD = CD0 + CDi | L/D = CL / CD
๐ Airfoil Aerodynamic Results
Configure inputs and click Compute to view results.
๐ 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 = 2ฯ (1 + 0.77 t/c) [ ฮฑ โ ฮฑ0L ] / [ 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.