High pressure turbine blade platform cooling and feed architecture

Thesis

In this thesis, novel cooling systems for gas turbine blade platforms were developed and assessed experimentally. A new highly modular facility was designed, manufactured, and commissioned to support investigations of a variety of different platform cooling systems. The facility is a high pressure, high temperature, transonic, linear cascade. A high degree of engine fidelity is achieved by: operating at scaled engine conditions; using test blades with full platform and fir tree root geometry; and including engine-representative internal hub seal structures and coolant feeds. The inclusion of hub seals, realistic coolant feeds, and full test blades allowed engine-feasible systems to be tested and representative leakage flows to be reproduced.

Two novel platform cooling systems were developed, using low-order numerical models, which utilise internal cooling features embedded directly into the platform surface. These systems are: a convective cooling passage design which draws coolant from the front seal, passes it through the platform, and exhausts into the rear seal; and an impingement jet array design which draws coolant from the shank cavity and similarly exhausts into the rear seal. Each system was integrated into cascade geometry rotor blades, manufactured in titanium via additive manufacturing, and used to experimentally determine metal effectiveness distributions across mainstream platform surfaces. Metal effectiveness distributions are determined from IR measurements of surface temperature. The cooling performance of each concept was quantified and compared to a baseline case with no cooling features, cooled solely by leakage flows.

The film cooling performance of front seal purge flow was investigated experimentally, to determine adiabatic film effectiveness, at a range of purge flow angles and mass ratios. The aerodynamic impact was assessed using a downstream area traverse system to determine total pressure loss coefficient distributions. A practical design to alter the purge flow angle in engine architectures was proposed and the viability of the system demonstrated using numerical modelling. A new experimental data processing technique was developed to correct adiabatic film effectiveness distributions determined from experiments with a non-uniform inlet temperature profile, such as that caused by a time-varying inlet thermal boundary layer.

Authors: Parker, JA

• Type of award: DPhil
• Level of award: Doctoral
• Awarding institution: University of Oxford
• Language: English
• Keywords: Linear Cascade; Turbine cooling; Turbomachinery; Experimental
• Subjects: Turbomachinery; Aerospace engineering