Stagnation-Point Thermal Load Estimation
This repository presents a computational model for estimating the wall temperature (Tₑ) at stagnation points on surfaces exposed to supersonic flow, based on an empirical boundary-layer formulation.
The model accounts for:
- Specific heat ratio (γ)
- Mach number
- Laminar flow recovery factor (r = 0.85)
Convective heating within the aerodynamic boundary layer is one of the primary sources of thermal load on exposed aerospace vehicle components.
To withstand elevated temperatures resulting from this phenomenon, insulating materials and ablative protection systems are typically employed in critical regions.
The stagnation-point temperature in the vicinity of surfaces directly facing the airflow can be estimated using the following empirical relation:
This simplified model assumes:
- Supersonic compressible flow
- Laminar boundary layer
- Ideal gas approximation
- No radiative heat transfer contribution
For a launch vehicle such as the VLS-1:
- Mach = 2
- Ambient air temperature = 0 °C (273 K)
Estimated nose fairing wall temperature:
≈ 185 °C
At Mach = 3:
Estimated wall temperature:
≈ 417 °C
These temperatures exceed the thermal tolerance of aluminum, which begins to lose significant mechanical strength at approximately 100 °C.
Therefore, stagnation regions — such as nose tips, leading edges, and aerodynamic protrusions — require:
- Thermal protection systems (TPS)
- Insulating coatings
- Ablative materials
- Advanced high-temperature alloys
- Preliminary aerothermal analysis
- Launch vehicle fairing design studies
- Thermal protection feasibility assessment
- Educational aerothermodynamics modeling
This project is licensed under the MIT License.
You are free to use, modify, and redistribute the code, provided proper attribution is maintained.
For collaboration inquiries, please connect via LinkedIn or email.