Abstract Over the past decades of preliminary aero-engine design a great effort has been invested in increasing steady-state efficiency to reduce missions fuel burn and thus CO2 emissions. Whilst pushing the performance cycle further to its limits previously minor deemed processes such as the transient behavior become more important because engines stability must still be granted for. It is therefore beneficial to integrate transient performance early into the overall aero-engine design tool chain to be able to predict the dynamic behavior from the beginning. This paper presents a methodology to couple predesign and transient performance simulation in order to get a more holistic picture of aero-engines early in the design process. This procedure entails extensive amount of data transfer throughout multi-disciplinary tools with different fidelity levels. This task is tackled using DLRs virtual engine platform GTlab (Gas Turbine Laboratory), which provides a geometric data model with abstract description of predesign components and standardized interfaces for data exchange. In order to demonstrate the proposed methodology a performance model of a turbofan similar to the V2500 aero-engine is used. For that purpose, a performance cycle is established providing boundary conditions for the preliminary aerodynamic engine design. The designed components provide necessary input data for the subsequent transient certification maneuver Eventually, parametric studies are conducted to show the impact of design variations on transient data such as the minimum surge margin and minimum tip clearance as well as on preliminary engine design.

Abstract The aim of the present paper is to improve the physical understanding of flow irregularities in the blade passing signal of turbomachinery rotors, since the novel stall warning method presented in part I is based upon those irregularities. For this purpose, a complementary instrumentation was used in a single stage axial compressor. A set of pressure transducers evenly distributed along the circumference surface mounted in the casing near the rotor tip leading edges is measuring the time-resolved wall pressures simultaneously to an array of transducers recording the chord-wise static pressures. The latter allows for plotting quasi-instantaneous 2D-pressure contours. Any occurring flow disturbances causing the before mentioned irregularity can later be classified using validated frequency analysis methods being applied to the data from the circumferential sensors. While leaving the flow coefficient constant, a continuously changing number of prestall flow disturbances appears to be causing the very spectral signature which is known from investigations on Rotating Instability. Any arising number of disturbances is matching a specific mode order to be found within the spectral signature. While the flow coefficient is reduced the propagation speed of prestall disturbances increases linearly as the speed seems to be independent from the clearance size. Data taken beyond the stalling limit demonstrate a complex superposition of stall cells and flow disturbances which the title “prestall disturbance” therefore doesn’t fit to precisely any more. Different convection speeds allow the phenomena to be clearly distinguished from each other.

Abstract This paper studies the effects of cut-corner in the recess vaned casing treatment (RVCT) on the performance of a low-speed axial fan. The solid casing fan and fans with corner-cut RVCT have been simulated respectively. The numerical simulation shows that the introduction of the RVCT and the modification of RVCT cut-corner size bring about obvious changes in fan performance. The RVCT establishes a connection between the RVCT channel and the main flow, contributing to the extra flow circulation in the tip region and resulting in an efficiency penalty for each RVCT configuration. The cut-corner is not necessarily beneficial to the improvement of the stable operating range. With the increase of cut-corner size at the rear wall of RVCT, the improvement of stall margin firstly demonstrates a drop to −8.82% before jumping to over 60%. RVCTs with different cut-corner sizes play different roles in the flow interaction between RVCT channel and main flow, and the stall margin poses a positive relationship with the mass flow passing through RVCT. The size of the cut-corner is responsible for the flow field in RVCT. The RVCT with small cut-corner size functions as a larger blade tip clearance and results in a smaller stall margin than the solid casing fan. Therefore, the size of the cut-corner in RVCT should be large enough to dredge the flow blockage in the tip region and delay the origin of stall.

Abstract Despite significant advancements in computational power and various numerical modeling in past decades, flow simulation of a multi-stage axial-flow compressor is still one of the most active areas of research, for it is the critical component in engine performance and operability, and there are so many elements that need to be looked into to predicting correct matching of the stages and accurate flow distribution inside the machine. Modeling unsteadiness, both deterministic and random types, and real geometries are among the most important features to be considered in such prediction. The authors have conducted in their previous studies a series of unsteady RANS (URANS) simulations of a 6.5-stage high-speed highly-loaded axial-flow compressor, and explored many unsteady effects as well as effects of real geometries such as Variable Stator Vane (VSV) clearance and inter-stage seal leakage flow on the compressor performance. However, all the analyses failed to predict correct stage matching, total pressure and temperature radial profiles, or mass-flow with adequate accuracies. In the present study, an Improved Delayed Detached Eddy Simulation (IDDES) with SST k-omega model is applied to the simulation of the same compressor configuration at aerodynamic design point. Fifth-order WENO scheme is employed for improved spatial accuracy to suppress significant increase in mesh size. Total number of mesh points are over 400 million for 1/10th sector model. Computations are ensemble averaged for 20 sector passage. Computed overall performance and flow field are compared with the compressor rig test data. The predictions of inter-stage total temperature radial profiles are noticeably improved over the URANS with the same mesh, discretization scheme and eddy turbulence model. Good comparison with the rig data indicates the current simulation is properly capturing the span-wise mixing phenomena. Unsteady flow field are compared between IDDES and URANS to locate the cause for the enhanced mixing. It is shown that components of Reynolds stress responsible for radial diffusion and anisotropic features are intensified in the tip leakage vortex at the rotor exit for the IDDES.

Abstract This paper discusses the effect of axial overlap on tip leakage flow behavior in a low speed tandem bladed axial compressor. Tandem bladed axial flow compressors provide an efficient way of obtaining a higher pressure rise with minimum number of stages as compared to a conventional design. The single stage tandem rotor compressor is designed to achieve a stage loading coefficient of 1.04 and flow coefficient of 0.9. A tip clearance of 3% (of span) is provided for the current study and it is maintained constant for all axial overlap configuration studies. A computational study of the designed stage was performed to investigate the behavior of tip leakage flow with different axial overlap and its effect on the overall performance of the stage. The application of tandem blading is only justified by prescribing high aerodynamic loading on both blades which increases the concern for associated tip leakage and corner separation losses. The study reveals the interaction of tip leakage flow from the fore and aft blade with the end wall flow and its influence on the main stream flow. The presence of an aft blade behind the fore seeks additional caution for optimum management of tip leakage flow. The leakage flows from the individual blades tend to roll up along with the shroud boundary layer resulting in the formation of momentum deficit fluid in the flow passage and at exit of rotors. The penetration of momentum deficit fluid in the main stream flow is directly related to the leakage flow structure from the fore and aft blade and also their interaction with the end wall flow. The axial overlap is seen to influence the leakage flow structure in two different ways. Firstly, it is observed that the initiation of tip leakage vortex of aft blade shifts from 30% to 10% of chord measured from leading edge, as axial overlap varies from 5% to −5% of chord in the tip region respectively. Secondly, a change in axial overlap also causes a change in passage geometry of the tandem blade interaction zone which has a first order influence on the mass flow rate through it. The accelerated flow through the passage geometry sustains high loading on the aft blade. It can also be utilized for mitigating the low momentum incoming fluid resulting from the mixing of tip leakage and end wall flow. The appropriate realization of these two effects has a potential to minimize the losses occurring due to tip leakage flow in the tandem compressor rotor and to achieve the benefit of higher loading with minimum number of stages.

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Abstract This paper aims at experimentally demonstrating the effects of axial slot casing treatment and tip gap variation on compressor performance, unsteady tip clearance flow, and stall inception features in a highly-loaded mixed-flow compressor at partspeed. Two tip gaps (0.32% and 0.64% of rotor blade chord at mid-span) were tested at three rotational speeds. A semicircular axial slot casing treatment improves compressor stability. The experimental results show that this casing treatment significantly moves the stability limit at partial speeds towards lower mass flow for both tip gaps, compared to the reference case without casing treatment. In the case of the compressor with casing treatment, efficiency increases for the large tip gap and decreases for the small tip gap. Dynamic pressure transducers installed in the casing upstream and along the rotor tip chord direction are used to detect the unsteady behavior of tip region flow and stall inception signals of the compressor. The characteristic frequency in the tip region decreases, and the oscillating amplitude first decreases and then increases during the throttling process, regardless of tip gap size or casing treatment. For axial compressors, by contrast, the observation in previous work has been an increase of the oscillating amplitude with decreasing flow coefficient. This is a surprising result of our work. Neither experiment nor CFD so far was able to explain why the trend in this mixed-flow compressor is different from the trend expected from axial compressors. The compressor stalls through the spike stall inception both with and without casing treatment. This observation also differs from recent studies on axial compressors, which demonstrated that casing treatments could change the type of stall inception. The unstable disturbance indicating initial stall inception initially appears in the blade tip region from blade mid-chord to trailing edge, and then propagates upstream towards the leading edge. This disturbance might be generated by the reversed flow separation near mid-chord.

Abstract NASA is investigating the potential of integrating acoustic liners into fan cases to reduce fan noise, while maintaining the fans aerodynamic performance. An experiment was conducted to quantify the aerodynamic impact of circumferentially grooved fan cases with integrated acoustic liners on a 1.5 pressure ratio turbofan rotor. In order to improve the ability to measure small performance changes, fan performance calculations were updated to include real gas effects including the effect of humidity. For all fan cases tested, the measured difference in fan isentropic efficiency was found to be less than the measurement repeatability for a torque-based efficiency calculation (≈ 0.2%), however, an unintended tip clearance difference between configurations makes it difficult to determine if circumferentially grooved fan cases degraded fan performance. Fan exit turbulence measurements showed a 1.5% reduction in total turbulence intensity between hardwall and circumferentially grooved fan cases in the tip vortex region, which is attributed to a disruption in the formation of the tip leakage vortex. This decrease in fan exit turbulence could potentially lead to a 1–2dB reduction in broadband rotor-stator interaction noise. Reduced aerodynamic performance losses associated with over-the-rotor liners could enable further fan noise reduction.

Abstract High pressure turbines are nowadays designed to a point where most design enhancements only yield marginal efficiency improvements. This challenges research facilities to reliably resolve ever smaller differences in efficiency caused by individual design changes. In recent years, immense efforts towards such highly accurate delta-efficiency measurements have been undertaken at the Large Scale Turbine Rig (LSTR). This paper comprises an overview of the applied methodology and the achievements on the basis of various validation cases. By thoroughly controlling the operation point and accounting for all variables affecting the efficiency η, the rig can resolve efficiency-differences Δη of ±0.1 % for a single day measurement. Four benchmark cases are investigated to validate the rig’s capabilities. First, the influence of tip clearance is investigated for a squealer-type geometry for swirling inflow. It is found that for an increase in tip clearance of 1 %, η is decreased by 2.68 %. Then, it is shown that a winglet-type tip geometry may improve the efficiency by Δη 0.33% in comparison to the squealer tip. Third, it is shown that these trends are similar for plain inflow, however swirl decreases efficiency by up to 1.25 % in comparison to plain inflow. Finally, the clocking-position of the combustor-module relative to the nozzle guide vanes is varied leading to efficiency differences of up to 0.52 %.

Abstract During the lifetime of a turbine stage, some of the blade tips may undergo changes due to mechanical rubbing with casing surface and also due to thermal oxidation. Understanding the effect these damaged blades have over the undamaged blades is essential to estimate the performance of the turbine stage in the operable tip clearance range. In this paper, the passage to passage aerodynamic interaction in a turbine stage is studied by modifying the tip gap of selected turbine blades and analyzing their effect on the neighboring blade passage flows. The experiments in this study are carried out in a single-stage low-speed axial turbine facility. All measurements are taken in the stationary frame of reference using a time-accurate differential dynamic pressure transducer mounted in a Kiel probe head. The experimental results from this study show that even with a significant increase on a selected blade’s tip clearance, its effect on the AFTRF turbine flow is only confined to its neighboring blade passage. The disturbances due to the altered tip clearance of one passage are not measurably propagated to its neighboring turbine passages. The changes made in one of the blades in a turbine stage do not significantly alter the aerodynamic performance of other blades. This result is particularly important for large-scale turbine research rigs such as AFTRF where the unsteady total pressure field is mapped in a time-efficient and phase-locked manner.

Abstract Long-term efforts have been made to understand loss generation and its reduction in the field of axial turbomachines. The traditional approach to losses for an isolated blade row considers the profile and the secondary losses as a result of viscous flow. The additional kinds of losses in the stage are connected with the shear stress in the mixing process. These losses result from the mixing of the main stream flow with 1) the stator leakage injected through the root axial gap and 2) the return of the tip leakage over the bucket shroud. This article focuses on the first type of mixing losses. The leakage to the main stream flow ratio and the root reaction are the two key parameters investigated in this study. The primary data source for this study is the experiment. An experimental single stage air turbine was modified to set and precisely measure the stator leakage flow. Three configurations of the single-stage test rig with different reaction levels were tested. The second data source for this study is CFD computation. These computations are applied to different geometries and conditions from the experiment; they are derived from real steam turbine stages designed in DSPW. The computations simulate multistage configuration and real steam is considered as the working fluid. CFD computations were performed in the commercial software ANSYS CFX. Each configuration task was computed in three iterative steps. Each step takes the distribution of the flow parameters on the boundary domains from the previous iteration. The final results from this ‘repeating boundary conditions’ approach better correspond with the real expansion in a multistage configuration. The two data sources are not directly comparable. The experiment is used for validation of the trends. The computations provide the possibility of a multi-parametric study. The multi-parametric study is necessary to obtain a more general loss model which can be used during turbine design. The evaluation of the experimental and numerical parts focuses on a comparison of the overall stage performance. Stage efficiency and reaction are presented in relation to the ratio between leakage and main stream flow.

Abstract In modern gas turbines, endwall contouring (EWC) is employed to modify the static pressure field downstream of the vanes and minimise the growth of secondary flow structures developed in the blade passage. Purge flow (or egress) from the upstream rim-seal interferes with the mainstream flow, adding to the loss generated in the rotor. Despite this, EWC is typically designed without consideration of mainstream-egress interactions. The performance gains offered by EWC can be reduced, or in the limit eliminated, when purge air is considered. In addition, EWC can result in a reduction in sealing effectiveness across the rim seal. Consequently, industry is pursuing a combined design approach that encompasses the rim-seal, seal-clearance profile and EWC on the rotor endwall. This paper presents the design of, and preliminary results from a new single-stage axial turbine facility developed to investigate the fundamental fluid dynamics of egress-mainstream flow interactions. To the authors’ knowledge this is the only test facility in the world capable of investigating the interaction effects between cavity flows, rim seals and EWC. The design of optical measurement capabilities for future studies, employing volumetric velocimetry and planar laser induced fluorescence are also presented. The fluid-dynamically scaled rig operates at benign pressures and temperatures suited to these techniques and is modular. The facility enables expedient interchange of EWC (integrated into the rotor bling), blade-fillet and rim-seals geometries. The measurements presented in this paper include: gas concentration effectiveness and swirl measurements on the stator wall and in the wheel-space core; pressure distributions around the nozzle guide vanes at three different spanwise locations; pitchwise static pressure distributions downstream of the nozzle guide vane at four axial locations on the stator platform.

Abstract The non-uniform disturbance in the circumferential direction is the main cause for the occurrence of rotating stall in turbomachinery. In order to study the effect of tip clearance leakage flow on rotating stall, the mixed-flow pump models with different tip clearances are numerically simulated, and then the energy performance curves and internal flow structures are obtained and compared. The results show that the computed pump efficiency and the internal flow field of the pump from numerical simulation are in good agreement with the experimental results. A saddle region appears in the energy performance curves of the three tip clearances, and with decrease in tip clearance, the head and efficiency of the mixed-flow pump increase and the critical stall point shifts, and the stable operating range of the mixed-flow pump decreases, which indicates that the mixed-flow pump stalls easily for smaller tip clearance. Under the deep stall condition, the influence of the leakage flow in the end wall area increases gradually with decrease in clearance. For small clearance, the leakage flow moves away from the suction surface to some distance to form a number of leakage vortex strips with the mainstream flow and flows over the leading edge of the next blade and then flows downstream into different flow passages, generating backflow and secondary flow separation at the blade inlet, which seriously damages the spatial structure of the inlet flow. This results in the earlier occurrence of stall. With increase in clearance, the leakage vortex develops along the radial direction towards the middle of the flow channel and large flow separation occurs in the downstream channel, which induces deep stall. For 0.8mm clearance, the whole impeller outlet passage is almost blocked by the backflow of the guide vane inlet, and a deep stall is induced.

Abstract In oilfield applications, an Electrical Submersible Pump (ESP) comprised of centrifugal pump, seal, and motor is placed inside a well to provide energy to lift reservoir fluids from the formation to the surface when the reservoir’s natural lift energy is insufficient. The use of an ESP as an Electrical Submersible Power Recovery Turbine (ESPRT) is the focus of this paper. As fluid is injected from the well surface to the pump discharge, through the pump stages, and out the pump intake, the ESPRT shaft rotates in reverse (counterclockwise from an uphole vantage point), compared to an ESP shaft. The moving fluid in the stages transmits energy to the impellers, which turns the motor shaft; the induction motor thereby becomes a generator. We simulated and optimized the performance of a conventional ESP stage using Computational Fluid Dynamics (CFD) and single-phase water. Before any laboratory or field testing, we simulated ESPRT performance at various speeds and flow rates to predict the performance of the turbine. The CFD prediction of the turbine follows the performance in the third quadrant of a typical four-quadrant pump curve. During the pre-field test CFD study, we did not attempt to predict the stage thrust. Later field trial results, however, showed the impellers operated in up thrust due to fluid flow through the turbine stages producing an upward reaction force on the impellers. We used the field test results to improve the CFD model and, afterwards, the stage design in order to reduce up thrust. The turbine, and pump upon which it is based, are designed to operate in down thrust. We investigated the effects of changing the diameter of the impeller balance holes and impeller-to-diffuser clearance at the impeller’s skirt and balance ring. Through several CFD iterations, we identified the optimum balance hole and clearance sizes; decreasing the size of the former feature and increasing the latter feature moved the impeller into the down thrust regime. Neither change reduced the stage head and corresponding power output.

Abstract The axial compressors of power-generation gas turbines have a high stage count, blades with low aspect ratios and relatively large clearances in the rear section. These features promote the development of strong secondary flows. An important outcome deriving from the convection of intense secondary flows is the enhanced span-wise transport of fluid properties mainly involving the rear stages, generally referred to as “radial mixing”. An incorrect prediction of this key phenomenon may result in inaccurate performance evaluation and could mislead the designers during the compressor design phase. As shown in a previous work, in the rear stages of an axial compressor the stream-wise vorticity associated with tip clearance flows is one of the main drivers of the overall span-wise transport phenomenon. Limiting it by circumferentially averaging the flow at row interfaces is the reason why a steady-state analysis strongly under-predicts radial mixing. To properly forecast the span-wise transport within the flow-path, an unsteady analysis should be adopted. However, due to the high blade count, this approach has a computational cost not yet suitable for industrial purposes. Currently, only the steady-state full-compressor simulation can fit in a lean industrial design chain and any model upgrade improving its radial mixing prediction would be highly beneficial for the daily design practice. To attain some progresses in RANS model, its inherent lack of convection of stream-wise vorticity must be addressed. This can be done by acting on another mixing driver, able to provide the same outcome, that is turbulent diffusion. In particular, by enhancing turbulent viscosity one can promote span-wise diffusion, thus improving the radial mixing prediction of the steady approach. In this paper, this strategy to update the RANS model and its application in simulations on a compressor of the Ansaldo Energia fleet is presented, together with the model tuning that has been performed using the results of unsteady simulations as the target.

Abstract A numerical study was conducted to evaluate the loss sensitivity of shrouded vs. unshrouded turbine rotor blades. Accuracy is demonstrated with a series of grid independence studies. Application of the methods is performed through various studies related to the effects of shrouding a High-Pressure Turbine (HPT) rotor blade for a NASA-specified N+3 timeframe single-aisle aircraft engine at takeoff conditions. Flat, Recessed, and Shrouded rotor configurations are evaluated at tip clearances from 0.25% to 4% of blade span. Mach # distributions, near-tip blade loading, and other flow characteristics are examined. Plots of stage efficiency vs. tip clearance are presented, with trends compared to available experimental data. It is shown that for the imposed boundary conditions, the addition of a shroud improves stage efficiency and significantly reduces sensitivity to tip clearance at higher clearance fractions. A casing recess is also shown to slightly increase sensitivity to tip clearance for tip clearances greater than 0.5%. Total pressure loss profiles vs. blade span are also compared, providing insight into the mechanisms behind the performance of the three configurations.

Abstract Erosion of compressor and turbine blades operating in extreme environment fouled with sand particles, ash or soot is a serious problem for gas turbine manufacturers and users. Indeed, operation of a gas turbine engine in such hostile conditions leads to drastic degradation of the aerodynamic performance of the components, mostly through surface roughness modification, tip clearance height increase or blunting of blade leading edges. To evaluate associated risks, the computation of particle trajectories and impacts through multiple turbomachinery stages by Computational Fluid Dynamics (CFD) seems a decent path but remains a challenge. The numerical prediction of complex turbulent flows in compressors and turbines is however necessary in such a context and validations are still required. Recently, Large-Eddy Simulation (LES) has shown promising results for compressor and turbine configurations for a wide range of operating conditions at an acceptable cost. With this in mind, this article presents the assessment of a LES solver able to treat turbomachine configurations to predict solid particle motion. To do so, the governing equations of particle dynamics are introduced using the Lagrangian formalism and are solved to compute locations and conditions of impact, namely particle velocity, angle and radius. The fully unsteady and coupled strategy is applied to blade geometries for studying the main areas and conditions of impacts obtained with LES. For comparison, a one-way coupling computation based on a mean steady flow field where only the Lagrangian particles are advanced in time is performed to evaluate the gain and drawbacks of both methods.

Abstract The concept of circumferential bending clearance based on Gauss Bimodal Function is proposed to suppress tip leakage flow (TLF) in a highly-loaded turbine cascade. In this method, a new vortex (BV) can be induced to mix with TLV in the middle of tip region and block the development of tip leakage vortex (TLV). Since the blocking effect divides the TLV into two parts, the tip leakage rate and loss of TLF can be reduced significantly. In order to reveal the mechanisms of blocking effect on leakage flow and its influencing factors, the research numerically investigates the effects of environmental conditions on the TLF development in a turbine cascade. The flow field analysis of the optimal bending clearance is in the first place, and then the effects of clearance heights (δ) and incidence angles (α) on the TLF characteristic and loss are investigated respectively. Results indicate that the blocking effect has a close relationship with the TLF characteristic, which can be divided into the BV migration, TLV-2 location and blocking loss. The nearer distance to the leading edge (LE) and farther distance to the suction side (SS) of BV means a less loss of TLF in bending clearance cases. The further distance away from blade tip and SS of TLV-2 means a larger-scale vortex with more loss. The additional loss in blocking region expands constantly with the increase of clearance height and incidence angle. The bending clearance has limited control effect on TLF with the variation of clearance height, especially the loss increases in Case 2%H. However, it has a strong adaptability with the change of incidence angle, the relative total pressure loss drops up to 16% in Case −5°.

Abstract The dielectric barrier discharge (DBD) plasma actuator, in which electrodes are asymmetric arranged, has already demonstrated its ability in flow control. In the present work, the configuration of multiple plasma actuators is placed at the suction side of the cascade top to realize the tip leakage control. However, massive configurations appear when the number of plasma actuators increases, resulting in the investigation of actuator configuration for tip leakage flow control becomes a challenge. The surrogate modelling approach provides a cheap and efficient method to investigate the effect of multiple plasma actuators on the tip leakage flow control. By constructing an approximation model, tip leakage mass flow rates of all configuration are obtained in the present work. What’s more, the flow structures in the tip clearance controlled by the plasma actuators are explained in the process of topological analysis. The results show that the tip leakage mass flow rate is decreasing with the number of active plasma actuators increasing. However, the decreasing would reach its limits in the process of adding plasma actuators. In the analysis of flow topology, single actuator would generate a small vortex at the suction side to cause an obstacle in the tip leakage flow. While the continuous arrangements of plasma actuator is beneficial to generate an induced vortex to diminish the tip leakage flow.

Abstract The pressure difference between suction and pressure sides of a turbine blade leads to the so-called phenomenon, the tip leakage flow, which most adversely affects the first-stage high-pressure (HP) turbine blade tip aerodynamics. In modern gas turbines, HP turbine blade tips are also exposed to extreme thermal conditions requiring the use of tip cooling. If the coolant jet directed into the blade tip gap cannot counter the leakage flow, it will simply add up to the pressure losses due to this leakage flow. Therefore, it is necessary to handle the design of tip cooling in such a way that the compromise between the aerodynamic loss and the gain in the tip cooling effectiveness is optimized. In this paper, the effect of tip cooling configuration on the turbine blade tip is investigated numerically both from the aerodynamics and thermal aspects in order to determine the optimum tip cooling configuration. The studies are carried out using the tip cross-section of General Electric E3 (Energy Efficient Engine) HP turbine first-stage blade for two different tip geometries, squealer tip and flat tip, where the number, location, and diameter of the cooling holes are varied. The study presents a discussion on the overall loss coefficient, the total pressure loss across the tip clearance, and the variation of heat transfer on the blade tip. The aerodynamic and heat transfer results are compared with the experimental data from literature. It is observed that increasing the coolant mass flow rate by using more holes or by increasing the hole diameter results in a decrease in the area-averaged Nusselt number on the tip floor, as expected. The findings show that both aerodynamic and thermal response of the squealer tips to the implementation of cooling holes is superior to their flat counterparts. Among the studied configurations, the squealer tip with larger number of cooling holes located towards the pressure side is highlighted as the configuration having the best cooling performance.