An experimental and numerical investigation of phantom cooling effects on cooled and uncooled rotating high pressure turbine blades in a full scale 1+1/2 stage turbine test is carried out. Objectives set to capture, separate, and quantify the effects of upstream vane film-cooling and leakage flows on the downstream rotor blade surface heat flux. Multiple series of 1+1/2 stage rotating high pressure turbine tests were carried out in the Air Force Research Laboratory, Turbine Research Facility, at Wright-Patterson Air Force Base, Ohio. A non-proprietary research turbine test article is uniquely instrumented with high frequency double-sided thin film heat flux gauges custom made at AFRL. High bandwidth, time resolved surface heat flux is measured on multiple film-cooled and non-film-cooled HPT rotor blades downstream of both film-cooled and non-film-cooled vane sectors. Upstream wake passing and heat flux is characterized on both rotor pressure and suction side surfaces, along with quantifying rotor phantom cooling effects from nonuniform 1st stage vane film cooling and leakage flows.
Fast response heat flux measurements quantify how rotor phantom cooling impacts the blade pressure side greatest; increasing along the pressure side towards the trailing edge. It is discovered upstream vane film-cooling alone can account for 50% of the rotor blade cooling effect, and even outweigh the rotor blade film cooling effect far from the blade showerhead holes. Added unsteady aero numerical simulation demonstrate how variations in inlet total temperature and incidence angle can also contribute to circumferentially non-uniform rotor heat flux.
Better understanding from this investigation aids modelling and design efforts in optimizing film cooling performance in real high pressure turbine flow fields. Understanding the behavior of such non-uniform circumferential rotor phantom cooling effects can be critical to optimize the efficiency, fuel consumption, range, and durability of advanced turbomachines.