Transient Material Response Analysis of Carbon-based Thermal Protection Systems for Rocket Nozzle Applications

The thermochemical erosion of carbon-based thermal pro- tection systems (TPS) needs to be accurately evaluated in order to get reliable performance predictions while designing solid rocket motors (SRM) and hybrid rocket engines (HRE). Moreover, it is important to ensure a proper sizing of the ablative TPS, leading to lightweight (i.e., minimum thickness) structures and preventing at the same time an excessive heating. In this context, reliable numerical models are required to correctly simulate the thermochemical and thermophysical behavior of ablative TPS. The aim of the present work is to perform a numerical investigation on nozzle ablation and its thermophysical material response. This is achieved by performing axisymmetric computational fluid dynamics (CFD) simulations including finite-rate ablative boundary conditions and pyrolysis gas species injection in the boundary layer, and simulations obtained with the recently developed Porous material Analysis Toolbox based on Open-FOAM (PATO). A specific numerical procedure for the generation of finite-rate thermochemical ablation tables for propulsive applications is derived. The results obtained by loosely coupling CFD and material response simulations employing the calculated finite-rate ablation tables are validated by comparison with firing tests data of a sub-scale Space Shuttle solid propellant booster employing a carbon-phenolic nozzle. The importance of accounting for finite-rate effects during the nozzle transient heating is highlighted.