Combined thermographic measurement and heat-flux compensation methods for aerodynamic heating evaluation in hypersonic flight

TL;DR

This study combines thermographic measurement and heat-flux compensation techniques with CFD validation to evaluate aerodynamic heating on a hypersonic projectile, providing detailed temperature and heat transfer insights during flight.

Contribution

It introduces a novel integrated method using high-speed thermography and CFD to accurately assess aerodynamic heating in hypersonic flight conditions.

Findings

Maximum surface temperature rise was 24.4 K above ambient.

Stanton number at stagnation point was 0.00366, consistent with CFD and empirical data.

Shock layer and temperature distribution were successfully visualized and validated.

Abstract

Novel thermographic measurement and heat-flux compensation methods combined for evaluating aerodynamic heating in hypersonic flight were developed using high-speed thermography. A hypersonic spherical projectile with a diameter of 8 mm was launched at approximately Mach 5 in the test section of a ballistic range. Shadowgraph imaging was conducted to visualize the flight trajectory and the shock layer. Thermographic measurement was performed using a high-speed infrared (IR) camera to obtain the surface temperature distribution of the projectile. The temperature distribution on the spherical surface was reconstructed from the thermographic data, by considering the photoresponse time of the photodetector of the IR camera and the geometric characteristics of the projectile trajectory. Furthermore, to validate the shock-layer geometry and aerodynamic heating characteristics, a computational fluid dynamics (CFD) simulation was also performed. The shadowgraph results showed that a detached shock wave and a shock layer were formed in front of the projectile, consistent with the CFD result. From the thermographic result, it was found that the maximum surface temperature rise during the flight was 24.4 K above the ambient temperature and it decreased with increasing distance from the stagnation point. The Stanton number distribution was estimated from the reconstructed surface temperature by assuming a one-dimensional transient heat conduction caused during the flight. The stagnation Stanton number was calculated to be 0.00366, which was also consistent with both the CFD result and a previously reported empirical correlation.