In this paper we establish a benchmark data set of a generic high-pressure turbine vane generated by direct numerical simulation (DNS) to resolve fully the flow. The test conditions for this case are a Reynolds number of 0.57 million and an exit Mach number of 0.9, which is representative of a modern transonic high-pressure turbine vane. In this study we first compare the simulation results with previously published experimental data. We then investigate how turbulence affects the surface flow physics and heat transfer. An analysis of the development of loss through the vane passage is also performed. The results indicate that free-stream turbulence tends to induce streaks within the near wall flow, which augment the surface heat transfer. Turbulent breakdown is observed over the late suction surface, and this occurs via the growth of two-dimensional Kelvin-Helmholtz spanwise roll-ups, which then develop into lambda vortices creating large local peaks in the surface heat transfer. Turbulent dissipation is found to significantly increase losses within the trailing-edge region of the vane.

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Direct Numerical Simulations of a High-Pressure Turbine Vane

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

Trends in turbomachinery turbulence treatments

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

A Novel Transient Technique to Determine Recovery Temperature, Heat Transfer Coefficient, and Film Cooling Effectiveness Simultaneously in a Transonic Turbine Cascade

This paper experimentally and numerically investigates the effects of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The base line turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75, and 1.01, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 910 5 , 1.0510 6 , and 1.510 6 based on a vane chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted a slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction sides of the vane, respectively. An increased Reynolds number was found to induce an earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and computational fluid dynamics predictions. DOI: 10.1115/1.2952381

Effects of Large Scale High Freestream Turbulence and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

This paper experimentally and numerically investigates the effect of large scale high freestream turbulence intensity and exit Reynolds number on the surface heat transfer distribution of a turbine vane in a 2-D linear cascade at realistic engine Mach numbers. A passive turbulence grid was used to generate a freestream turbulence level of 16% and integral length scale normalized by the vane pitch of 0.23 at the cascade inlet. The baseline turbulence level and integral length scale normalized by the vane pitch at the cascade inlet were measured to be 2% and 0.05, respectively. Surface heat transfer measurements were made at the midspan of the vane using thin film gauges. Experiments were performed at exit Mach numbers of 0.55, 0.75 and 1.01 which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 9 × 105 , 1.05 × 106 , and 1.5 × 106 , based on true chord. The experimental results showed that the large scale high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the vane as compared to the low freestream turbulence case and promoted slightly earlier boundary layer transition on the suction surface for exit Mach 0.55 and 0.75. At nominal conditions, exit Mach 0.75, average heat transfer augmentations of 52% and 25% were observed on the pressure and suction side of the vane, respectively. An increased Reynolds number was found to induce earlier boundary layer transition on the vane suction surface and to increase heat transfer levels on the suction and pressure surfaces. On the suction side, the boundary layer transition length was also found to be affected by increase changes in Reynolds number. The experimental results also compared well with analytical correlations and CFD predictions.Copyright © 2007 by ASME

Hong Wu

Papers

Effects of Large Scale High Freestream Turbulence and Exit Reynolds Number on Turbine Vane Heat Transfer in a Transonic Cascade

Effects of large scale high freestream turbulence, and exit reynolds number on turbine vane heat transfer in a transonic cascade