โ๏ธ Lift & Airfoil Calculator
Analyze NACA 4-digit airfoils using thin airfoil theory with finite-wing corrections. Outputs CL, CD, CM, polar diagram, and airfoil shape.
โก Fortran 90 Engine
Double Precision (IEEE 754)
โ ISO / ASME Validated
Fluid Mechanics
๐ Solver Telemetry
โ ACTIVE
๐๏ธ Views
123
โก Solves
97
๐พ Downloads
549
๐ฆ Fortran Code
4.5 KB
๐
Released
Jun 2026
โฑ๏ธ Latency
< 1 ms
๐ชฝ Airfoil Shape
๐ Configuration
Key Equations:
Thin airfoil: CL = 2ฯ(ฮฑ โ ฮฑL0)
Finite wing: CL3D = CL2D / (1 + 2ฯ/(ฯ AR))
Induced drag: CDi = CLยฒ/(ฯeAR)
NACA 4-digit camber and thickness distributions.
Thin airfoil: CL = 2ฯ(ฮฑ โ ฮฑL0)
Finite wing: CL3D = CL2D / (1 + 2ฯ/(ฯ AR))
Induced drag: CDi = CLยฒ/(ฯeAR)
NACA 4-digit camber and thickness distributions.
๐ Results & Polar Diagrams
Configure inputs and click Analyze to view results.
๐ Calculation Methodology
Mathematical Model
Thin airfoil theory provides CL = 2ฯ(ฮฑโฮฑL0) for 2D. Finite-wing corrections use lifting-line theory. Induced drag follows from the Oswald span efficiency factor. Post-stall CL reduction uses a simplified Viterna-like model.
NACA 4-Digit Series
The first digit is max camber (% chord), the second is position of max camber (tenths of chord), and the last two digits give max thickness (% chord). The thickness and camber distributions define the airfoil shape.
Assumptions
- Inviscid thin airfoil theory + viscous drag estimate.
- Parasitic drag from turbulent flat plate correlation with form factor.
- Simplified post-stall behavior.
- No compressibility correction (low Mach).
- Rectangular planform assumed for AR.