Effects of slot bleed injection over a contoured end wall on nozzle guide vane cooling performance: Part II - Thermal measurements

Endwall heat transfer has become a major issue in the design of the inlet nozzle guide vane region of modern gas turbine engines. To compensate for high rates of convective heat transfer and the uncertain flow pattern along endwall surfaces, coolant flows are often excessive and distributed in a less than optimum fashion. In many instances, coolant is carried away or mixed into the core flow by the secondary flows without being effective. There is a need for more effective cooling concepts. In this paper, the results of an experimental study examining the thermal performance of bleed injection through an inclined slot positioned upstream of the nozzle airfoil leading edge plane are presented. This paper demonstrates that this type of combustor bleed cooling is a promising cooling concept. Testing is performed in a large-scale, guide vane cascade comprised of three airfoils between one contoured and one flat endwall. The Reynolds number, based upon approach velocity and true chord length, is 350,000 and the approach flow is with large-scale, high-intensity (9.5%) turbulence. Combustor bleed cooling flow is injected ahead of a contoured endwall with bleed-to-core mass flow ratios as high as 6%. Measurements are taken to document core flow temperature distributions at several axial positions within the cascade to evaluate surface adiabatic effectiveness values and local heat transfer coefficients. This film cooling arrangement offers significant thermal protection. The coolant is shown to provide thermal protection over most of the endwall as well as portions of the pressure and suction surfaces of the airfoils. To achieve this coverage, combustor bleed flow must be strong enough to overcome the influence of endwall region secondary flows.