This paper investigates the influence of coolant injection on the aerodynamic and thermal performance of a rotor blade cascade with endwall film cooling. A 7 blade cascade of a high-pressure-rotor stage of a real gas turbine has been tested in a low speed wind tunnel for linear cascades. Coolant is injected through ten cylindrical holes distributed along the blade pressure side. Tests have been preliminarily carried out at low Mach number (Ma2is = 0.3). Coolant-to-mainstream mass flow ratio has been varied in a range of values corresponding to inlet blowing ratios M1 = 0–4.0. Secondary flows have been surveyed by traversing a 5-hole miniaturized aerodynamic probe in two downstream planes. Local and overall mixed-out secondary loss coefficient and vorticity distributions have been calculated from measured data. The thermal behaviour has been also analysed by using Thermochromic Liquid Crystals technique, so to obtain film cooling effectiveness distributions. All this information, including overall loss production for variable injection conditions, allow to draw a comprehensive picture of the aero-thermal flow field in the endwall region of a high pressure rotor blade cascade.

This content is only available via PDF.
You do not currently have access to this content.