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Efficiency improvements for gas turbines are strongly coupled with increasing turbine inlet temperatures. This imposes new challenges for designers for efficient and adequate cooling of turbine components. Modern gas turbines use bleed air from the compressor to inject into the stator/rotor rim seal cavity to prevent hot gas ingestion from the main flow, while cooling the rotor disk. This purge flow interacts with the main flow field and static pressure field imposed by the turbine blades. The complex interaction causes nonuniform and jet-like penetration of the purge flow into the main flow field and therefore affects the endwall heat transfer on the rotor. In order to improve the understanding of purge flow effects on rotor hub endwall heat transfer, measurements in the 1.5-stage axial turbine facility LISA at ETH Zurich have been performed. An unshrouded, high-pressure representative turbine design with 3D blading and extended endwall contouring of the rotor into the cavity seal has been tested. A state-of-the-art measurement setup with a high-speed infrared camera and thermally managed rotor insert has been used to perform high-resolution heat transfer measurements on the rotor. Three different purge flow rates were investigated with regard to hub endwall heat transfer. Additionally, steady-state CFD simulations were performed to complement the experiments. It was found that the local heat transfer rate changes up to ±20% depending on the purge flow rate. The main part of the purged air is ejected at the endwall trough location and swept towards the rotor suction side caused by the interaction of main flow and the cavity extended endwall design. The presence of low momentum purge flow is locally reducing the heat transfer rate. Changes in adiabatic wall temperature and heat transfer depending on purge rate are observed from the platform start up to the cross passage migration of the secondary flow structures. INTRODUCTION Over the past years, various design approaches were developed to increase efficiency and reduce the losses due to secondary flows in turbomachines. Up to 30-50% of the total aerodynamic losses can be attributed to secondary flow losses according to (Sharma and Butler, 1987). (Langston, 2001) provides a condensed review about secondary flows in turbomachines. A promising approach to reduce and control the cross passage flow due to the pressure gradient is endwall contouring. Studies by (Harvey et al., 1999) showed that the usage of non-axisymmetric endwall contouring is a powerful tool for reducing secondary flows with comparable effects to lean and skew of blades. Cascade measurements confirmed a reduction of secondary losses by 30% (Hartland et al., 1999). An efficiency increase of 1.0% and 0.4% is also reported by (Schuepbach et al., 2008) and (Snedden et al., 2009) respectively.

Purge Flow Effects on Rotor Hub Endwall Heat Transfer with Extended EndwallContouring into the Disk Cavity