Vasudevan Kanjirakkad

Flow in a Rotating Cavity With Axial Throughflow at Engine Representative Conditions

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.

A New Experimental Facility to Investigate Combustor–Turbine Interactions in Gas Turbines With Multiple Can Combustors

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.

Aerodynamic and aerothermal investigation of the flow around an HPT rotor shroud: PIV measurements

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.

Passive Shroud Cooling Concepts for HP Turbines: CFD Based Design Approach

One option to improve the cycle efficiency of current state-of-the-art aero engines is to increase the turbine inlet temperature. Since this temperature is above the melting temperature for the alloys utilised in the turbine component already today, efficient cooling methods must be developed that consider both aerodynamic and aerothermal aspects of cooling. Here, the goal is to extract as little as possible secondary air from the main hot gas cycle for cooling and to use this coolant then aerodynamically and aerothermally as efficient as possible. The paper to be presented documents a CFD based design approach that lead to a new passive shroud cooling concept and the definition of its operational parameters. By using a simple one dimensional method [10] for predicting the aerodynamic losses resulting from such a cooling configuration in connection with 3d Navier-Stokes solvers (RANS) for predicting film cooling effectiveness contours on the rotor shroud surfaces, the new cooling configuration was developed. The concept was then tested and confirmed experimentally as documented in more detail in Part 2 of this paper. It is noted that only 70% of the coolant mass flow required for the current configuration was used for the new concept whereas the aerodynamic efficiency measured remained nearly constant. Improving upon existing passive shroud cooling systems where the coolant is injected directly into the labyrinth of the shroud, the new approach comprises cooling holes that inject the coolant upstream of the labyrinth and through the stator platform into the main passage flow. Here, it is important that the bulk of the coolant is placed below the dividing streamlines between main passage flow and labyrinth flow. Thereby, it can be achieved that the major part of the coolant indeed reaches the thermally loaded target surfaces on the shroud bottom at various axial gaps due to different operating points of the turbine. Besides the improved film-cooling effectiveness measured, the second important aspect of the new concept is the achievement of as small as possible additional aerodynamic losses due to coolant ejection into a high speed flow region. It will be shown that both goals can be achieved by the new concept. Furthermore, CFD results on film-cooling performance and aerodynamic losses will be shown and compared with experimental data.

Aerodynamic and aerothermal investigation of the flow around an HPT rotor shroud: heat transfer and cooling effectiveness

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 describe the aerothermal performance of a passive shroud cooling strategy using measured heat transfer and adiabatic cooling effectiveness data. Improved physical understanding of the shroud aerodynamics as reported in the companion paper Lehmann et al. [1] will be used here to support the analysis. Highly resolved experimental heat transfer data was acquired on the shroud, the fins and on the shroud underside with the thin heater film method. The distribution of the adiabatic cooling effectiveness on the rotor shroud was measured with a combination of the Ammonia-Diazo and a foreign gas sampling technique. The measurements are complemented by steady numerical computations of the turbine stage. Due to the impact of vortical flow structures in the over shroud cavities, the Nusselt numbers on the shroud top surfaces were found to be of the same order as on the shroud underside. The passive shroud cooling concept was found to provide quite efficient and uniform cooling to the over-shroud surfaces while the distribution of coolant on the shroud underside was significantly affected by the rotor secondary flow.

Improving the fidelity of aerodynamic probes using additive manufacturing

Purpose This paper aims to offer the aerodynamic testing community a new procedure for manufacturing high-quality aerodynamic probes suitable for 3D flow measurements with consistent geometry and calibration by taking advantage of the additive manufacturing technology. Design/methodology/approach The design methodology combines the advantages and flexibilities of computer aided design (CAD)/computer aided manufacturing (CAM) along with the use of computational fluid dynamics to design and analyse suitable probe shapes prior to manufacturing via rapid prototyping. Findings A viable procedure to design and possibly batch manufacture geometrically accurate pneumatic probes with consistent calibration is shown to be possible through this work. Multi-jet modelling prototyping methods with wax-based support materials are found to be a cost-effective method when clean and long sub-millimetre pressure channels are to be cut. Originality/value Utilisation of the geometry consistency that is made possible by 3D printing technology for the design and development of pneumatic probes is described. It is suggested that the technique could lead to batch production of identical probes, thus avoiding precious time of a skilled labourer and elaborate individual calibration requirement.

A New Experimental Facility to Investigate Combustor-Turbine Interactions in Gas Turbines With Multiple Can Combustors

A New Experimental Facility to Investigate Combustor-Turbine Interactions in Gas Turbines With Multiple Can Combustors

Leading Edge Shielding Concept in Gas Turbines With Can Combustors

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and, hence, in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels. The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so-called showerhead region. In the past, investigators have tested many different showerhead configurations by varying the number of rows, inclination angle, and shape of the cooling holes. However, the current leading edge cooling strategies using showerheads have not been shown to allow a further increase in turbine temperature without the excessive use of coolant air. Therefore, new cooling strategies for the first vane have to be explored. In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e., to protect, the first vane leading edges from the high heat loads. In this way, the stagnation region at the leading edge and the showerhead of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs. This paper numerically analyzes the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and nonrestrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.