This review addresses recent literature on turbine passage aerodynamics and endwall heat transfer; articles that describe the endwall flow and cooling problems are summarized, recent activity on improving endwall aerothermal design is discussed, improved cooling schemes are proposed, and methods for managing secondary flows to allow more effective cooling are suggested. Much attention is given to aerodynamic losses associated with secondary flows developed near the endwalls. The endwall region flowfield is influenced by the stagnation zones established as the endwall approach flow boundary layer meets the airfoil leading edges, by the curvature of the passages, by the steps and gaps on the endwall surface ahead of and within the passage, by the leakage and coolant flows introduced through the endwall surface ahead of and within the passage, by the tip leakage flows between the blades and shroud in the rotor endwall region, and by many more effects. Recent combustor redesigns have flattened the turbine inlet temperature profile and have raised the turbine inlet temperatures. This, coupled with a continued need to improve engine durability and availability, has spurred strong interest in thermal control of the turbine endwall regions. Thus, much of the literature presented herein is focused on endwall cooling and, in particular, the effects of near-endwall secondary flows on endwall cooling.

Turbine Endwall Aerodynamics and Heat Transfer

Aerodynamic losses due to the formation of a leakage vortex near the tip section of rotor blades form a significant part of viscous losses in axial flow turbines. The leakage flow, mainly induced by the pressure differential between the pressure side and suction side of a rotor tip section, usually rolls into a streamwise vortical structure near the suction side part of the blade tip. The current study uses the concept of a tip platform extension that is a very short “winglet” obtained by slightly extending the tip platform in the tangential direction. The use of a pressure side tip extension can significantly affect the local aerodynamic field by weakening the leakage vortex structure. Phase averaged, time accurate total pressure measurements downstream of a single stage turbine facility are provided from a total pressure probe that has a time response of 150 kHz. Complete total pressure maps in all of the 29 rotor exit planes are measured accurately. Various pressure and suction side extension configurations are compared against a baseline case. The current investigation performed in the Axial Flow Turbine Research Facility (AFTRF) of the Pennsylvania State University shows that significant total to total efficiency gain is possible by the use of tip platform extensions.Copyright © 2001 by ASME

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Aerodynamic Tip Desensitization of an Axial Turbine Rotor Using Tip Platform Extensions

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability The flow at combustor exit can therefore have high residual swirl A good deal of research concerning the flow within the combustor is available in open literature The impact of swirl on the aerodynamic and heat transfer characteristics of a HP turbine stage is not well understood, however A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern At the turbine inlet plane, yaw and pitch angles of over +/-40 degrees have been simulated The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the non-dimensional parameters for aerodynamics and heat transfer are matched to engine conditions The research turbine was the unshrouded MT1 design By design, the centre of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2 For the current investigation, the clocking position was such that the vortex centre was aligned with the vane leading edge (every second vane) Both the aligned vane and the adjacent vane were characterised This paper presents measurements of HP vane surface and endwall heat transfer for the two vane positions The results are compared with measurements conducted without swirl The vane surface pressure distributions are also presented The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra The aerodynamic and heat transfer characterisation presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors The findings not only have implications for the vane aerodynamic design, but also for the cooling system designCopyright © 2011 by Rolls-Royce plc

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

This paper is focused on the effect of film hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film cooling holes are divided into four rows and loosely distributed on the platform. Four film cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film cooling effectiveness distributions on the platform was obtained using pressure sensitive paint (PSP) technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.Copyright © 2008 by ASME

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

Non-axisymmetric endwall profiling is a promising method to reduce secondary losses in axial turbines. However, in high-pressure turbines, a small amount of air is ejected at the hub rim seal to prevent the ingestion of hot gases into the cavity between the stator and the rotor disk. This rim seal purge flow has a strong influence on the development of the hub secondary flow structures. This paper presents time-resolved experimental and computational data for a one-and-1/2-stage high work axial turbine showing the influence of purge flow on the performance of two different non-axisymmetric endwalls and the axisymmetric baseline case. The experimental total-to-total efficiency assessment reveals that the non-axisymmetric endwalls lose some of their benefit relative to the baseline case when purge is increased. The first endwall design loses 50% of the efficiency improvement seen with low suction, while the second endwall design exhibits a 34% deterioration. The time-resolved computations show that the rotor dominates the static pressure field at rim seal exit when purge flow is present. Therefore, the purge flow establishes itself as jets emerging at the blade suction side corner. The jet strength is modulated by the first vane pressure field. The jets introduce circumferential vorticity as they enter the annulus. As the injected fluid is turned around the rotor leading edge a streamwise vortex component is created. The dominating leakage vortex has the same sense of rotation as the rotor hub passage vortex. The first endwall design causes the strongest circumferential variation in the rim seal exit static pressure field. Therefore, the jets are stronger with this geometry and introduce more vorticity than the other two cases. As a consequence the experimental data at rotor exit shows the greatest unsteadiness within the rotor hub passage with the first endwall design.Copyright © 2009 by ASME

Influence of Rim Seal Purge Flow on the Performance of an Endwall-Profiled Axial Turbine

An important problem that arises in the design and the performance of axial flow turbines is the understanding, analysis, prediction and control of secondary flows. Sieverding1 has given a review of secondary flow literature, covering up to 1985. In this paper a brief review of pre-1985 work is given, and then a survey of open literature secondary flow investigations since the Sieverding review is presented. Most of the studies reviewed deal with plane or annular cascade flows. Tip clearance effects are not covered. The basic secondary flow picture for a turbine cascade, as measured and verified by a number of investigators is described. Recent work that shows refined secondary flow vortex structures is examined. A flow parameter based on inlet boundary layer properties used to predict horseshoe vortex swirl is presented. Work on secondary flow loss reduction, involving airfoil geometry, endwall fences and endwall contouring is briefly reviewed. A new leading edge bulb geometry that has demonstrated impressive loss reduction is considered. It is concluded that accurate routine prediction of secondary flow losses has not yet been achieved, and must await either a better turbulence model or more experiments to reveal new endwall loss production mechanisms. Lastly, loss is examined from the standpoint of entropy generation.

L.S. Langston

Papers

Secondary Flows in Axial Turbines—A Review