To prevent melting of turbine blades, numerous cooling schemes have been developed to cool the blades using cooler air from the compressor. Unfortunately, the clearance gap between adjacent hub sections allows coolant to leak into the hub region. Coolant flow also leaks into the hub region through gaps between individual stages. The results of a combined experimental and computational study of cooling along the hub of a first stage turbine blade caused by leakage flows are discussed in detail. Additionally, this study examines a novel cooling feature, called a microcircuit, which combines internal convective cooling with external film cooling. For the experimental investigation, scaled up blades were tested in a low speed wind tunnel. Adiabatic effectiveness measurements were made with infrared thermography of the entire hub region for a range of leakage flow conditions. For the computations, a commercially available computational fluid dynamics (CFD) code, FLUENT 6.0, was used to simulate the various flows. Results show that featherseal leakage flows provide small cooling benefits to the hub. Increases in featherseal flow provide no additional cooling to the hub region. Unlike the featherseal, leakage flows from the front rim provide ample cooling to the hub region, especially the leading edge of the blade passage. None of the leakage flows provide significant cooling to the pressure side region of the hub or trailing edge suction side. With the addition of the hub microcircuits, there is improved hub cooling of the suction side of the blades. Though the coolant exit uniformity was low and affected by the featherseal flow, the microcircuits were shown to provide more cooling along the hub region. Good agreements were observed between the computational and experimental results, though computations over-predicted front rim cooling and microcircuit uniformity. / Master of Science

Further improvements of the thermal efficiency of gas turbine cycle are closely coupled to the increase of turbine inlet temperature. This requires intensive and efficient cooling of the blades. In this perspective, experimental investigations of the gas temperature and heat transfer distribution around the airfoil are of primary importance. The present work aims at the development of two measurement techniques based on applications of the thin film sensors: the two-layer gauge for the wall heat transfer determination and the dual thin film probe for flow temperature measurements. Both techniques are used in short duration tunnels of the von Karman Institute (VKI) under engine representative conditions and are able to resolve both time-averaged component and time-resolved component i.e. periodic blade passing events at ~5-7 kHz with harmonics up to 50 kHz. In order to derive the wall heat flux with the two-layer gauge, the unsteady conduction equation is solved in the two-layer substrate using the measured value of the wall temperature as a boundary condition. The gauges are extensively calibrated and the data reduction method is validated on a blade of the second stator of the VKI turbine. A very good repeatability is achieved. Measurements are also performed on the complex geometry of a blade tip in a cascade configuration revealing the high three dimensionality of the flow. The dual thin film probe combines the operation of two thin films and determines the flow temperature from two independent heat flux measurements. The probe is calibrated and then validated with measurements downstream a cascade. The robustness and the reliability of the probe are also demonstrated by measurements downstream of the rotor and the second stator of the VKI turbine.