Budimir Rosic

Control of Shroud Leakage Loss by Reducing Circumferential Mixing

Shroud leakage flow undergoes little change in the tangential velocity as it passes over the shroud. Mixing due to the difference in tangential velocity between the main stream flow and the leakage flow creates a significant proportion of the total loss associated with shroud leakage flow. The unturned leakage flow also causes negative incidence and intensifies the secondary flows in the downstream blade row. This paper describes the experimental results of a concept to turn the rotor shroud leakage flow in the direction of the main blade passage flow in order to reduce the aerodynamic mixing losses. A three-stage air model turbine with low aspect ratio blading was used in this study. A series of different stationary turning vane geometries placed into the rotor shroud exit cavity downstream of each rotor blade row was tested. A significant improvement in flow angle and loss in the downstream stator blade rows was measured together with an increase in turbine brake efficiency of 0.4 %.

The Importance of Shroud Leakage Modeling in Multistage Turbine Flow Calculations

Three-dimensional steady multistage calculations, using the mixing plane approach, are compared with experimental measurement in a low-speed three-stage model turbine. The comparisons are made with two levels of shroud seal clearance, one representative of a real turbine and one with minimal seal clearance and almost no shroud leakage. Three different calculations are compared. The first computes the main blade path with no modeling of shroud leakage. The second includes a simple model of shroud leakage using sources and sinks on the end-walls, and the third is a multiblock calculation with all leakage paths and cavities computed. It is found that neglect of shroud leakage makes the computed velocity profiles and loss distributions significantly different to those measured. Simple modeling of shroud leakage gives some improvement but full calculation of the leakage flows and cavities is necessary to obtain good agreement between calculation and measurement.

The Influence of Shroud and Cavity Geometry on Turbine Performance: An Experimental and Computational Study— Part II: Exit Cavity Geometry

Imperfections in the turbine annulus geometry, caused by the presence of the shroud and associated cavity, have a significant influence on the aerodynamics of the main passage flow path. In this paper, the datum shroud geometry, representative of steam turbine industrial practice, was systematically varied and numerically tested. The study was carried out using a three-dimensional multiblock solver, which modeled the flow in a 1.5 stage turbine. The following geometry parameters were varied: inlet and exit cavity length, shroud overhang upstream of the rotor leading edge and downstream of the trailing edge, shroud thickness for fixed casing geometry and shroud cavity depth, and shroud cavity depth for the fixed shroud thickness. The aim of this study was to investigate the influence of the above geometric modifications on mainstream aerodynamics and to obtain a map of the possible turbine efficiency changes caused by different shroud geometries. The paper then focuses on the influence of different leakage flow fractions on the mainstream aerodynamics. This work highlighted the main mechanisms through which leakage flow affects the mainstream flow and how the two interact for different geometrical variations and leakage flow mass fractions.

Blade Lean and Shroud Leakage Flows in Low Aspect Ratio Turbines

Blade lean, i.e., nonradial blade stacking, has been intensively used over the past in the design process of low aspect ratio gas and steam turbines. Although its influence on turbine efficiency is not completely understood, it has been proved as an effective way of controlling blade loading and secondary flows on blade passage endwalls. Three-dimensional blade designs in modern industrial practice are usually carried out using clean endwalls. The influence of the leakage flows on three-dimensional blade design is traditionally neglected. This paper presents an experimental study where two different stator blades, with different levels of compound lean, were tested in a low speed three-stage model turbine with the shroud leakage flow geometry representative of industrial practice. The experimental measurements were compared with numerical tests, conducted on the same blade geometries. The influence of the compound lean on the stator flow field was analyzed in detail. In order to analyze the combined effects of both the stator hub and the rotor shroud leakage flow on the blade lean, in the second part of the paper a numerical study on a two stage turbine with both leakage flow paths representative of a real turbine was carried out. Performance of three different stator blade designs (two different levels of compound lean and a straight blade) was investigated. The aim of this study is to understand the mechanism and the consequence of the stator blade lean on stage performance in an environment with leakage flows and associated cavities.

Controlling Tip Leakage Flow Over a Shrouded Turbine Rotor Using an Air-Curtain

This paper describes an experimental and computational investigation into the performance of an air-curtain seal used to control the leakage flow over the tip shroud of a turbine rotor. The results show that a seal of this type has the potential to reduce or eliminate shroud leakage whilst having a practical level of clearance between the stationary and moving components. The experimental measurements were undertaken using a single-stage low-speed air turbine equipped with a continuous circumferential nozzle in the casing to deliver an axisymmetric jet into the cavity over the rotor shroud. The jet was angled at 45° to the axial direction so that its momentum opposed the shroud leakage flow. In this arrangement the air-curtain was able to sustain the pressure difference between the inlet and outlet of the rotor blade row without any leakage. The test facility had comprehensive instrumentation for obtaining accurate measurements of turbine efficiency that were corrected for the externally supplied additional flow required for the air-curtain. Measurements were obtained for a range of jet flows and show the change in the turbine efficiency as the jet flow is increased. The measurements have been compared with calculations.Copyright

Integrated Combustor and Vane Concept in Gas Turbines

This paper numerically investigates the interaction between multiple can combustors and the first vane in an industrial gas turbine with 16 can combustors and 32 vanes in order to find ways of reducing the overall cooling requirements. Two promising concepts for the overall cooling reduction are presented. In the first, by minimising the axial distance between the combustor wall and the vane, the stagnation region at the LE of every second vane can be effectively shielded from the hot mainstream gases. The LE shielding allows continuous cooling slots to be used (as an alternative to discrete cooling holes) to cool downstream parts of the vane using a portion of the saved LE showerhead cooling air. The second concept proposes a full combustor and first vane integration. In this novel concept the number of vanes is halved and the combustor walls are used to assist the flow turning. All remaining vanes are fully integrated into the combustor walls. In this way the total wetted area of the integrated system is reduced, and by shielding the LEs of the remaining vanes the total amount of cooling air can be reduced. The proposed combustor and first vane integration does not detrimentally affect the aerodynamics of the combustor and vane system. The concept also simplifies the design and should lower the manufacturing costs.Copyright

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.

Influence of Large Wake Disturbances Shed From the Combustor Wall on the Leading Edge Film Cooling

The first vane leading edge film cooling is challenging because of the highest thermal load and the complex flow interaction between the hot mainstream gas and the coolant flow. This interaction varies significantly from the stagnation region to the regions of high curvature and acceleration further downstream. Additionally, in industrial gas turbines with multiple combustor chambers around the annulus the first vane leading edges may also be exposed to large wake disturbances shed from the upstream combustor walls.The influence of these vortical structures on the first vane leading edge film cooling is numerically analysed in this paper. In order to assess the capabilities of the flow solver TBLOCK to simulate these complex interactions an experimental test case is modelled numerically. The test case is available in the open literature and consists of a cylindrical leading edge and two rows of film cooling holes representative of industrial practice. A LES turbulence modelling strategy with WALE sub-grid scale (SGS) model is applied and compared against experimental results. Based on this validation it is decided to analyse also the wake–leading edge interaction, dominated by large scale unsteady vortical structures, using the same WALE sub-grid scale LES model. The initial flow domain with the cylindrical leading edge and cooling holes is extended to incorporate the effect of the combustor wall, which is modelled as a flat plate with a square trailing edge. The location and the size of the plate are scaled to be representative of industrial practice: the plate is located upstream from the leading edge at a distance twice the leading edge diameter, and the thickness of the plate is one half of the leading edge diameter. Two different clockwise positions of the vertical combustor wall model were investigated and compared with the datum configuration: the former where the axis of the plate and the leading edge are aligned (central wake location), the latter with the combustor wall circumferentially shifted up by a quarter of the leading edge diameter (circumferentially shifted wake location).Numerical predictions show that the shed vortices from the combustor wall trailing edge have a highly detrimental effect on the leading edge film cooling by periodically removing the coolant flow from the leading edge surface. This results in an increased unsteady thermal load. These negative effects are less significant in the case of circumferentially shifted wake, due to the combined action of both shed vortices.Copyright

Uncertainty Quantification of Leakages in a Multistage Simulation and Comparison With Experiments

April 10-15, 2016 Abstract The present paper presents a numerical study of the impact of tip gap uncertainties in a multistage turbine. It is well known that the rotor gap can change the gas turbine efficiency but the impact of the random variation of the clearance height has not been investigated before. In this paper the radial seals clearance of a datum shroud geometry, representative of steam turbine industrial practice, was systematically varied and numerically tested. By using a Non-Intrusive Uncertainty Quantification simulation based on a Sparse Arbitrary Moment Based Approach, it is possible to predict the radial distribution of uncertainty in stagnation pressure and yaw angle at the exit of the turbine blades. This work shows that the impact of gap uncertainties propagates radially from the tip towards the hub of the turbine and the complete span is affected by a variation of the rotor tip gap. This amplification of the uncertainty is mainly due to the low aspect ratio of the turbine and a similar behavior is expected in high pressure turbines. .

Development and Aerothermal Investigation of Integrated Combustor Vane Concept

This paper presents the development and aerothermal investigation of the Integrated Combustor Vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010 [1], the conventional Nozzle Guide Vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept is developed using the inhouse CFD code TBLOCK. Aerothermal experiments are conducted using a modular high speed linear cascade, designed to model the flow at the combustor-vane interface. The facility comprises two can combustor transition ducts and either four Conventional Vanes (CVs) or two Integrated Vanes (IVs). The experimental study validates the linear CFD-simulations of the IV development. Annular full stage CFD-simulations, used to evaluate aerodynamics, heat transfer and stage efficiency, confirm the trends of the linear numerical and experimental results and thus demonstrate the concept’s potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile- and endwall-loss. Combined with an improved rotor performance these aerodynamic benefits result in a gain in stage efficiency of above 1%, illustrated by unsteady stage simulations. A distinct reduction in HTC levels on vane surfaces, in the order of 25%–50%, and endwalls is observed and attributed to an altered state of boundary layer thickness. The development of IV’s endwall- and LE-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.Copyright