The prediction of compressor drum cavity heat transfer is an important factor in the overall design of an aero engine. The rotationally dominated flow field within the cavity governs the heat transfer conditions by suppressing the motion of the fluid. Without heating, the fluid in the outer region of the cavity can approach solid body rotation. The outer cavity fluid is disturbed by the bore flow at the inner radius. The resultant bore flow vortex has been shown to exhibit many different modes of behaviour, dependent on the Rossby number. At higher Rossby number the bore flow vortex has been shown to break down into a precessing radial arm. It has also been shown that the hot drive arm (shroud) between the compressor stages destabilises the flow field through natural convection. This paper presents data from the Sussex Multiple Cavity Rig, which matches the fluid dynamic conditions of a compressor bore in terms of axial throughflow, rotational Reynolds number and Grashof number. It features titanium alloy discs, which are instrumented with surface thermocouples. This paper presents data which helps to separate the effects of throughflow Reynolds number, rotational Reynolds number and Grashof number on the dimensionless disc temperature profiles. In order to illustrate the flow structures this paper presents a hybrid RANS/LES model for the two highest Reynolds number cases. For these cases, the numerical simulations show a change from stable to unstable stratification with an increase in the bore to shroud temperature ratio in good qualitative agreement with the measured data.
This paper describes a new modular experimental facility that was purpose-built to investigate flow interactions between the combustor and first stage nozzle guide vanes (NGVs) of heavy duty power generation gas turbines with multiple can combustors. The first stage turbine NGV is subjected to the highest thermal loads of all turbine components and therefore consumes a proportionally large amount of cooling air that contributes detrimentally to the stage and cycle efficiency. It has become necessary to devise novel cooling concepts that can substantially reduce the coolant air requirement but still allow the turbine to maintain its aerothermal performance. The present work aims to aid this objective by the design and commissioning of a high-speed linear cascade, which consists of two can combustor transition ducts and four first stage NGVs. This is a modular nonreactive air test platform with engine realistic geometries (gas path and near gas path), cooling system, and boundary conditions (inlet swirl, turbulence level, and boundary layer). The paper presents the various design aspects of the high pressure (HP) blow down type facility, and the initial results from a wide range of aerodynamic and heat transfer measurements under highly engine realistic conditions.
The flow and heat transfer inside HP compressor rotating cavities are buoyancy driven and are known to be extremely difficult to predict. The experimental data of Laser-Doppler Anemometry (LDA) measurements inside an engine representative cavity rig is presented in this paper. Traverses using a two component LDA system have been carried out in the shaft bore and the cavity regions in order to map the axial and tangential velocity components. The velocity data is collected for a range of Rossby, Rotational and Axial Reynolds numbers, Ro, Reθ and Rez: 0.08 < Ro < 0.64, 7 × 105 < Reθ < 2.83 × 106, 1.2 × 104 < Rez < 4.8 × 104 and for values of the buoyancy parameter βΔT, 0.284 < βΔT < 0.55. Numerical study using unsteady Reynolds Averaged Navier-Stokes (URANS) simulations have been carried out to elucidate flow details for a few selected cases. The experimental results revealed that the Swirl number (Xk) varies from a value < 1 near the bore to near solid body rotation at increased radii within the cavity. The analysis of frequency spectrum of the tangential velocity inside the cavities has also shown the existence of pairs of rotating and contra-rotating vortices. There is generally satisfactory agreement between measurements and CFD simulations. There is also convincing evidence of two or more separate regions in the flow dominated by the bore flow and rotation.
An experimental study has been conducted to investigate the aerothermal performance of a shrouded high pressure turbine blade in a large scale rotating rig. The rotor blade and the associated shroud and casing geometry have been modelled in a large scale low speed turbine rig that was designed to investigate a novel passive shroud cooling methodology. The objective of the present paper is to provide a detailed description of the flow field around the rotor blade shroud. The improved physical understanding of the shroud flow gained from this study will be used to analyse the aerothermal performance of the shroud cooling strategy as reported in a companion paper, Lehmann et al. [1]. Experiments have been carried out using endoscopic PIV to identify and understand salient flow features that exist upstream and downstream of the shroud as well as within the shroud cavities. The measurements are complemented by steady and unsteady numerical computations of the turbine stage. The study identifies the existence of important vortical structures within the shroud cavities that not only interact with the main passage flow but also modify the amount and distribution of the shroud leakage flow in a manner that has major implications for shroud cooling and heat transfer. A detailed shroud flow model is derived and used to elucidate the causes and consequences of the flow pattern observed. The model emphasises the circumferentially asymmetric nature of the cavity flow structures caused by the presence of the inter shroud gap that in turn influences the production, interaction and dissipation of such vortical structures.
This paper describes a new modular experimental facility that was purpose-built to investigate flow interactions between the combustor and first stage nozzle guide vanes of heavy duty power generation gas turbines with multiple can combustors. The first stage turbine nozzle guide vane is subjected to the highest thermal loads of all turbine components and therefore consumes a proportionally large amount of cooling air that contributes detrimentally to the stage and cycle efficiency. It has become necessary to devise novel cooling concepts that can substantially reduce the coolant air requirement but still allow the turbine to maintain its aerothermal performance. The present work aims to aid this objective by the design and commissioning of a high-speed linear cascade which consists of two can combustor transition ducts and four first stage nozzle guide vanes. This is a modular non-reactive air test platform with engine realistic geometries (gas path and near gas path), cooling system, and boundary conditions (inlet swirl, turbulence level and boundary layer). The paper presents the various design aspects of the high pressure blow down type facility, and the initial results from a wide range of aerodynamic and heat transfer measurements under highly engine realistic conditions.
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