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

Heat transfer: a review of 2002 literature

This paper presents experimental measurements and computational predictions of surface and end wall heat transfer for a high-pressure (HP) nozzle guide vane operating as part of a full HP turbine stage in an annular rotating turbine facility, with and without inlet temperature distortion (hot streaks). A detailed aerodynamic survey of the vane surface is also presented. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough, UK. This is a short-duration facility, which simulates engine-representative M, Re, nondimensional speed, and gas-to-wall temperature ratio at the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation combustor simulator, capable of simulating well-defined, aggressive temperature profiles in both the radial and circumferential directions. This work forms part of the pan-European research program, TATEF II. Measurements of HP vane and end wall heat transfer obtained with inlet temperature distortion are compared with results for uniform inlet conditions. Steady and unsteady computational fluid dynamics (CFD) predictions have also been conducted on vane and end wall surfaces using the Rolls-Royce CFD code HYDRA to complement the analysis of experimental results. The heat transfer measurements presented in this paper are the first of their kind in that the temperature distortion is representative of an extreme cycle point, and was simulated with good periodicity and with well-defined boundary conditions in the test turbine.

Effect of Simulated Combustor Temperature Nonuniformity on HP Vane and End Wall Heat Transfer: An Experimental and Computational Investigation

An inclined turbulent jet discharging a passive scalar into a turbulent crossflow is investigated under conditions of favorable, zero and adverse streamwise pressure gradient. Experiments are conducted in water by means of magnetic resonance velocimetry and magnetic resonance concentration measurements. The velocity ratio and density ratio are equal to one for all cases. The flow configuration is relevant to film cooling technology, the molecular properties of the fluid being immaterial in the fully turbulent regime. Under favorable pressure gradient (FPG), the streamwise acceleration tilts the jet trajectory toward the wall, which would be beneficial for the film cooling performance. However, the counter-rotating vortex pair is strengthened in the accelerating flow by streamwise stretching. Also, the crossflow boundary layer is significantly thickened by increasingly adverse pressure gradient, which affects the mass transfer from the jet. Overall, the more intense counter-rotating vortices and the thinner boundary layer associated with increasingly FPG enhance the scalar dispersion into the main flow, hampering the film cooling performance in terms of surface effectiveness.

R.W. Radomsky

Papers

Elevated freestream turbulence effects on heat transfer for a gas turbine vane