Passage Vortex Research Articles

Vortex Patterns Investigation and Enstrophy Analysis in a Small Scale S-CO2 Axial Turbine

Supercritical carbon dioxide (S-CO2) Brayton cycle system is a promising closed-loop energy conversion system frequently mentioned in the automotive and power generation field in recent years. To develop a suitable design methodology for S-CO2 turbines with better performance, an understanding of the vortex flow patterns and associated aerodynamic loss inside a S-CO2 turbine is essential. In this paper, a hundred-kilowatt level S-CO2 axial turbine is designed and investigated using a three-dimensional transient viscous flow simulation. The NIST Span and Wagner equation of state model that considers the real gas effects is utilized to estimate the thermodynamic properties of the supercritical fluid. The numerical methods are experimentally validated. The results indicates that the aspect ratio and tip-to-hub ratio are different in the S-CO2 turbine from that in the gas turbine, and the vortex flow patterns are influenced notably by these geometrical parameters. Both the vortex structure and moving tracks of passage vortices are changed as a result of large centrifugal force. An interaction between tip leakage vortex and hub passage vortex is observed in the impeller passage and its formation and development mechanism are revealed. To further explore the aerodynamic loss mechanism caused by vortex interaction, the energy loss in the impeller passage is analyzed with the enstrophy dissipation method, which can not only accurately calculate the energy loss but also estimate how the vortical motions occur. It is found that the enstrophy and energy loss can be effectively reduced by vortex interaction between tip leakage vortex and hub passage vortex. The results in this study would increase the knowledge of vortex flow patterns in S-CO2 turbine and the proposed enstrophy production method can be used intuitively to provide a reference for flow vortical motion study in turbines.

Combined flow control with full-span slot and end-wall boundary layer suction in a large-camber compressor cascade

To investigate combined flow control strategies with full-span slot and end-wall (EW) boundary layer suction, a large-camber compressor cascade was treated with a full-span slot configuration firstly. Then, two combined flow control schemes with different EW boundary layer suction slots were set up based on the full-span slot configuration. Finally, the performance of the baseline cascade and the slotted cascades was evaluated and compared numerically. The results show that the full-span slotted scheme can eliminate the trailing edge (TE) separation near the blade midspan and weakens the corner separation. However, the EW secondary flow still develops on the blade suction surface (SS) before the full-span slot outlet and separates to form the wall vortex (WV). In contrast, the two combined flow control schemes can effectively inhibit the development of the EW secondary flow on the blade SS, contributing to the smaller total pressure loss and stronger static pressure recovery capacity. Between them, the combined flow control scheme with full-chord-length EW suction slot behaves better. It can effectively eliminate the concentrated shedding vortex, WV and corner vortex in the cascade passage, and the passage vortex is also weakened to some extent. Therefore, the TE separation and the corner separation are basically eliminated. The total pressure loss, flow turning angle and static pressure coefficient can be increased by an average of −38.4%, 3.1° and 16.2% in the combined flow control cascade with full-chord-length EW suction slot respectively, which is beneficial for improving the stability margin and static pressure recovery capacity of compressor blades.

Effect of the bionic chamber position on the aerodynamic performance in a transonic compressor cascade

A bionic chamber structure has been investigated in a linear compressor cascade to suppress corner separation of suction side, reduce flow loss, and improve cascade performance. Effects of the bionic chamber at various positions on flow field structures and aerodynamic performance have been studied by numerical simulation. The investigation illustrates that the bionic chamber can increase turbulence kinetic energy near the endwall. Additionally, corner separation along the pitchwise direction is suppressed, and an interaction of concentrated shedding vortex and passage vortex is weakened. Meanwhile, the existence of the bionic chamber is able to diminish the flow loss and enhance the diffusion ability. With the increase of the number of the bionic chamber from the leading edge to the trailing edge, the range of low-energy fluid shrinks more in the pitchwise direction and expands more apparently in the spanwise direction. The bionic chamber arranged in the range of 25% to 50% chord length produces the best effect, which can reduce total pressure loss coefficient by 9.43% and increase static pressure coefficient by 8.35%.

Near Tip Loss Control With a Winglet Baffle Cavity Tip in a Turbine Cascade

Abstract The aerodynamic performance of a winglet baffle cavity tip is investigated at different inlet incidences from −12.5 to +12.5 deg. This blade tip shows a geometry feature with a pressure side winglet and a baffle within the tip cavity. The experimental studies were carried out in a large-scale linear cascade. Numerical methods were also used to obtain the detail flow physics. The baffle on the tip divides the cavity vortex into two main parts, and enhances flow mixing over the tip. As the tip leakage flow exits the tip near the baffle and the cavity corner, flow separation occurs over the suction side and its blockage effect reduces the local tip leakage mass flow rate. The additional pressure side winglet reduces the contraction coefficient on the pressure side squealer. Compared with the squealer tip, the winglet baffle cavity tip reduces the tip leakage mass flow by 12.1% and the near tip loss by 4.2%. Removing the baffle on the tip also results in an increase of the loss. As the incidence of the incoming flow decreases, the loss near the tip reduces mainly due to a reduction of the strength of the passage vortex, which develops from the casing endwall. At all incidences studied, the winglet baffle cavity tip shows better aerodynamic performance than the squealer tip.

Unsteady flows of a highly loaded turbine blade with flat endwall and contoured endwall

Strong secondary flow results in substantial aerodynamic loss for highly-loaded turbine. Accurate prediction of the complicated flow structures proposes challenges to the widely used Reynolds Average Navier-Stokes (RANS) approach. This work employs the hybrid RANS/Large Eddy Simulation (LES) method to study the unsteady flows for a highly-loaded turbine blade, with both flat endwall and optimized contoured endwall. Evolution of the unsteady flows in the endwall region is analyzed, with emphasis on the loss generation mechanism. Results show that for the flat endwall, the horseshoe vortex system is highly unsteady and is a significant source of unsteadiness in both the passage and the wake region. It contributes to the unsteady passage vortex, also the earlier breakdown of the trailing edge shedding vortex. The Probability Density Function (PDF) histogram of velocity in the wake region is bimodal, implying the perturbations from two mechanisms. For the contoured endwall, the unsteady evolution of the horseshoe vortex is blocked, which results in significantly reduced unsteadiness, also the merge between the horseshoe vortex with the passage vortex is prevented. Effect of the flow unsteadiness on the loss generation is assessed based on the entropy generation rates contributed by the time-averaged flow and the fluctuations. The effect of unsteadiness is two-fold: unsteady perturbations trigger the vortex breakdown into small-scale structures and thus weakened wake velocity deficit and loss generation; however, in both the passage and the wake region the entropy generation by the fluctuations are remarkable. For the contoured endwall, due to the reduced flow unsteadiness, loss generation contributed by the fluctuations are much smaller compared to the flat endwall. The results highlight the importance of including the loss generation by the fluctuations, also a possible mechanism to reduce the secondary loss by attenuating the flow unsteadiness with the contoured endwall.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 1: Aerodynamics

Abstract The first-stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics is not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage. The test section involves an engine-representative combustor–turbine interface geometry, combustor coolant flow, and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 2: Thermal Measurements

Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage (NGV). The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing, and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

Numerical Predictions of Turbine Cascade Secondary Flows and Heat Transfer With Inflow Turbulence

Abstract Turbine passage secondary flows are studied for a large rounded leading edge airfoil geometry considered in the experimental investigation of Varty et al. (“Vane Suction Surface Heat Transfer in Regions of Secondary Flows: The Influence of Turbulence Level, Reynolds Number, and the Endwall Boundary Condition,” ASME J. Turbomach., 140(2), p. 021010) using high resolution large eddy simulation. The complex nature of secondary flow formation and evolution are affected by the approach boundary layer characteristics, components of pressure gradients tangent and normal to the passage flow, surface curvature, and inflow turbulence. This paper presents a detailed description of the secondary flows and heat transfer in a linear vane cascade at exit chord Reynolds number of 5 × 105 at low and high inflow turbulence. Initial flow turning at the leading edge of the inlet boundary layer leads to a pair of counter-rotating flow circulation in each half of the cross plane that drives the evolution of the pressure side and suction side of the near-wall vortices such as the horseshoe and leading edge corner vortex. The passage vortex for the current large leading edge vane is formed by the amplification of the initially formed circulation closer to the pressure side (PPC) which strengthens and merges with other vortex systems while moving towards the suction side. The predicted suction surface heat transfer shows good agreement with the measurements and properly captures the augmented heat transfer due to the formation and lateral spreading of the secondary flows towards the vane midspan downstream of the vane passage. Effects of various components of the secondary flows on the endwall and vane heat transfer are discussed in detail.

Large-Eddy Simulation of Low-Pressure Turbine Cascade with Unsteady Wakes

To better understand the wake effects at low Reynolds numbers, large-eddy simulations of a 50% reaction low-pressure turbine stage and a linear cascade with two different bar wake generators were carried out for a chord Reynolds number of 50,000. For the chosen front-loaded high-lift airfoil, the endwall structures are stronger than for more traditional mid-loaded moderate-lift airfoils. By comparing the 50% reaction stage results with the bar wake generator results, insight is gained into the effect of the three-dimensional wake components on the downstream flow field.For the cases with bar wake generator, the endwall boundary layer is growing faster because of the relative motion of the endwall with respect to the freestream. The half-width of the wake is approximately matched for the larger one of the two considered bar wake generators. To improve the quality of the phase-averaged flow fields, the proper orthogonal decomposition was employed as a filter to remove the low-energy unsteady flow field content. Both the mean flow and filtered phase-averaged flow fields were analyzed in detail. Visualizations of the phase-averaged flow field reveal a periodic suppression of the laminar suction side separation from the downstream airfoil even for the smaller bar wake generator. The passage vortex is entirely suppressed for the 50% reaction stage and for the larger bar wake generator. Furthermore, the phase-averaged data for the 50% reaction stage reveal a new longitudinal flow structure that is traced back to near-wall wake vorticity. This flow structure is missing for the bar wake generator cases.

Influence of tip clearance distribution on blade vibration of vaneless radial turbine

As vehicle turbochargers are developed toward higher performance and lower turbo lag, high cycle fatigue (HCF) of radial turbine blades is becoming increasingly common which greatly threatens the reliability of turbochargers. Tip leakage vortex is one of potential sources of blade excitation and it’s profoundly influenced by blade tip clearance. This paper studies the influence of tip clearance distribution on blade excitation of a vaneless radial turbine via experimentally validated one-way fluid-structure interaction (FSI) numerical method. The results suggest that blade vibration response is significantly influenced by tip clearance distribution in the meridional direction. Generalized energy method is proposed to determine the key factors for blade excitation. The results manifest that complex distributions of harmonic pressure amplitude on the blade dominate blade vibration response. Detailed flow field analysis is carried out to further investigate the mechanism of blade excitation. The results show that distributions of harmonic pressure amplitude on pressure surface (PS) and suction surface (SS) are both dominated by tip leakage vortex, whereas the roles that tip leakage vortex plays are quite different. Specifically, tip leakage vortex influences harmonic pressure amplitude on SS directly because of short distance between vortex core and SS, whereas it influences harmonic pressure amplitude on PS indirectly by interfering the evolution of passage vortex. This research can guide new designs for durable vaneless radial turbines without sacrificing aerodynamic performance.

Vortex‐Dominated Aeolian Activity at InSight's Landing Site, Part 1: Multi‐Instrument Observations, Analysis, and Implications

AbstractWe report the aeolian changes observed in situ by NASA's InSight lander during the first 400 sols of operations: Granule creep, saltation, dust removal, and the formation of dark surface tracks. Aeolian changes are infrequent and sporadic. However, on sols, when they do occur, they consistently appear between noon to 3 p.m., and are associated with the passage of convective vortices during periods of high vortex activity. Aeolian changes are more frequent at elevated locations, such as the top surfaces of rocks and lander footpads. InSight observed these changes using, for the first time, simultaneous in‐situ and orbital imaging and high‐frequency meteorological, seismological, and magnetic measurements. Seismometer measurements of ground acceleration constrain the timing and trajectory of convective vortex encounters, linking surface changes to source vortices. Magnetometer measurements show perturbations in magnetic field strength during the passage of convective vortices consistent with charged‐particle motion. Detachment of sand‐scale particles occurs when high background winds and vortex‐induced turbulence provide a peak surface friction wind speed above the classic saltation fluid threshold. However, detachment of dust‐ and granule‐scale particles also occurred when the surface friction wind speed remained below this threshold. This may be explained by local enhancement of the surface roughness and other effects described here and further studied in Part 2 (Baker et al., 2021). The lack of saltation and bright dust‐coated surfaces at the InSight landing site implies surface stability and the onset of particle motion may be suppressed by dust “cushioning.” This differentiates the InSight landing site from other areas on Mars that exhibit more aeolian activity.

Performance enhancing for a highly loaded tandem cascade by endwall incoming vortex–corner separation interaction

In highly loaded tandem compressor cascades, corner separations can still exist. In order to eliminate corner separations in highly loaded tandem compressor cascades, incoming vortex–corner separation interaction mechanism was investigated. Different schemes of the vortex generators, which located at different pitchwise locations and could generate vortexes with different rotation directions, were designed and investigated numerically. Results show that, severe corner separation occurred at the front blade passage of the tandem cascade; by utilizing flow control method of incoming vortex–corner separation interaction, the corner separation could be reduced significantly. The optimal control effect of incoming vortex on corner separation was achieved with anticlockwise rotation and the vortex generator is located right ahead of the leading edge of tandem cascade, a maximum loss coefficient reduction of 21.8% being achieved. Different from single blade configuration, the boundary layer of tandem cascade was regenerated at rear blade suction surface due to the injection flow from blade gap between the two blades. Though corner separations could be reduced at both conditions, the loss of tandem cascade with clockwise incoming vortex is higher than that with anticlockwise vortex, and a smaller corner separation region at suction surface was achieved by utilizing clockwise vortex. The mechanism was that anticlockwise incoming vortex reduced the corner separation but increased secondary flow, while clockwise vortex enhanced passage vortex and decreased secondary flow. For clockwise incoming vortex near pressure surface, the vortex would be divided into two parts at the leading edge of rear blade, one would go through the blade gap and deteriorate flow fluid near rear blade suction surface, the other flowed downstream along pressure surface. The rotation direction of different incoming vortexes became the same as the passage vortex at rear blade passage of tandem cascade, which was mainly due to the effect of secondary flow.

Investigation of the flowfield on a vaneless counter-rotating turbine with seal cavity flow ejection

The sealing of the rotor-rotor gap and rotor disk cooling are vital to the safe operation of the vaneless counter-rotating turbine(VCRT). In order to quantifies the influence of the wheel-space cavity flow on the VCRT aerodynamic performance, and to improve turbine efficiency of the VCRT at certain rim seal ejection rates, numerical studies which considered the effects of rotor-rotor rim seal flow ejection are carried out in this paper. The three dimensional unsteady computational fluid dynamic analysis of a VCRT at the engine conditions are performed, and the seal flow ejected downstream of the high pressure rotor row at six sealing flow rate are modeled. The interaction among the high pressure rotor trailing shock wave, the downstream secondary flow and the seal flow has been studied and quantitatively characterized as a function of the purge ejection rate. Numerical results show that seal flow- mainstream flow interaction is entirely dominated by the high pressure rotor trailing edge shock at the hub, low pressure hub passage vortex and the mixing of the sealing flow from wheelspace and mainstream. When the mass flow rate of the coolant is smaller than some threshold value, the shock loss of the high pressure rotor and hub secondary flow loss of the low pressure rotor are decreased with the increasing of the coolant mass flow rate. It causes that the VCRT efficiency is gradually increased. On condition that the amount of the seal flows is beyond the threshold value, the key roles in modification of the VCRT performance are changed. The increment of the hub secondary flow loss and the mixing loss are gradually larger than the decrement of the shock loss. As a result, the turbine efficiency gradually decreases.

Structure and Topology Analysis of Separated Vortex in Forward-Swept Blade

Flow separation commonly affects the stability of turbomachines, especially under low-flowrate conditions. Compared with conventional blades, a forward-swept blade is more efficient at high flowrates. However, experiments and numerical simulations show that a forward-swept blade produces an unstable region under low flowrate. In this paper, the topological analysis is used to analyze the structure and size of flow separation in forward swept blades. Three-dimensional structure and formation mechanism of vortices in forward-swept blades are analyzed using the cross-section flow pattern method. For forward-swept blades, flow separation mainly occurs at the blade tip and corner, accompanied by clear velocity fluctuations, the break-up of shed vortices, and diffusion. With decreasing flowrate, the shedding vortices move forward and the speed of vortex annihilation gradually decreases. In addition, the number of singularities in the rotor passage increases with the decrease of flow rate, and the region affected by shedding vortex increases. The rotating direction of internal vortex in turbomachinery is fixed. The pressure surface, passage vortex, and concentrated shedding vortex were found to rotate clockwise, whereas the suction surface, corner vortex, and shedding vortex rotate in a counterclockwise direction.

Comparison of Endwall Loss Reduction Techniques in a High-Lift Turbine Passage

Abstract The development of techniques that reduce the losses in the endwall region is an important area of research as it supports an increase in the turbine design space through the use of higher lift blade designs while maintaining high efficiency. Several active and passive shape contouring methods that reduce losses generated by the secondary flow have been developed and investigated on a high-lift front-loaded low-pressure turbine research profile in a low-speed linear cascade configuration. This paper summarizes and compares alterations to the three-dimensional secondary flow field by the application of three different techniques: blade profile contouring, optimized endwall shape contouring, and localized low mass coefficient jets. Each method was applied to identical research blade profiles and compared in the same linear cascade wind tunnel resulting in a unique perspective into the loss reduction mechanisms associated with each technique. The design strategy will be discussed along with a detailed description of changes to the secondary flow field using in- and out-of-passage total pressure loss measurements and high-speed stereoscopic particle image velocimetry. The key findings include the loss reduction mechanisms associated with each approach and the manipulation of key endwall flow structures such as the passage vortex and a strong suction surface corner separation.

Unsteady Steam Turbine Optimization Using High-Fidelity Computational Fluid Dynamics

Abstract This paper presents the computational methodology and experimental investigations accomplished to enhance the efficiency of a turbine stage by applying non-axisymmetric profiling on the rotor hub wall. The experimental setup was a two-stage axial turbine, which was tested at “LISA” test facility at ETH Zurich. The first stage was considered to create the flow history for the second stage, which was the target of the optimization. The hub cavity of the second stage was designed with large dimensions as a requirement of a steam turbine. The goal was to optimize the interaction of the cavity leakage flow with the rotor passage flow to reduce the losses and increase efficiency. The computational optimization was completed using a genetic algorithm coupled with an artificial neural network on the second stage of the test turbine. Unsteady time-accurate simulations were performed using in-house developed “MULTI3” solver. Besides implementing all geometrical details (such as hub and tip cavities and fully 3D blade geometries) from the experimental setup into the computational model, it was learned that the unsteady upstream effect could not be neglected. A novel approach was introduced using unsteady inlet boundary conditions to consider the multistage effect while reducing the computational cost to half. The importance of this implementation was tested by performing a steady simulation on the optimized geometry. The predicted efficiency gain from steady simulations was 4.5 times smaller (and negligible) compared to the unsteady approach. Excluding the cavity geometry was also assessed in a different simulation setup showing 3.9% over-prediction in the absolute efficiency value. Comprehensive steady and unsteady measurements were performed utilizing pneumatic, fast response aerodynamic probe (FRAP), and fast response entropy probe (FENT) on the baseline and profiled test cases. The end wall profiling was found to be successful in weakening the strength of the hub passage vortex by a 19% reduction in the under-over turning. As a result, the blockage was reduced near the hub region leading to more uniform mass flow distribution along the span. The flow angle deviations at the higher span position were also corrected due to better control of the flow angles. Furthermore, the improvements were confirmed by reductions in entropy, secondary kinetic energy, and pressure unsteadiness. The accurate computational implementations led to an excellent agreement between the predicted and measured efficiency gain.

Assessment of the Influence of Boundary Layer Ingestion (BLI) on the Axial Fan

Boundary Layer Ingestion (BLI) is a technology which is effective to reduce fuel consumption of future commercial aircrafts. However, the inlet distortion caused by BLI has a negative on the engine performance. This paper investigates the effect of the inlet total pressure distortion on the aerodynamic performance of the fan rotor. Unsteady numerical analysis was conducted on JAXA TechClean Fan with circumferential distortion caused by BLI. As the result, under the high distorted condition, the peak efficiency of the rotor decreases by 0.622 points. It is found that the main factor of the efficiency reduction is tip-wall passage vortex.

Pulsed Suction Control in a Highly Loaded Compressor Cascade With Low Suction Flowrates

Abstract The influence of pulsed suction (PS) on flow separation and aerodynamic performance in a highly loaded compressor cascade is experimentally studied herein. The excitation frequency is investigated as it determines the effectiveness of PS in flow control. Low suction flowrates are examined to analyze the potential of PS in providing a satisfactory cascade performance. For comparison, the corresponding parameters of steady continuous suction (SCS) are studied as well. Oil-flow visualizations and steady and unsteady pressure data are used to characterize the control effects of SCS and PS. The experimental results validate the efficacy of PS in controlling flow separation, even at a reduced suction flowrate of 0.1%. It suppresses passage vortex (PV), improving aerodynamic performance. PS provides a better control effect than SCS at different excitation parameters, which can be attributed to twofold main reasons: first, at the same suction flowrate, PS has a larger suction momentum than SCS during the suction phase, resulting in a stronger suction force and having a more profound effect on the flow characteristics; and second, owing to the introduction of pulsed excitation to the suction, PS creates additional vortex structures that energize the boundary layer by transporting high momentum freestream fluid near the wall. PS is also effective at a higher incidence angle, but its control effect is reduced.

Investigation and Application of Bio-Inspired Microscale Structures to Turbine Endwall for Enhancing Film Cooling Performance

Abstract In this paper, inspired by micro-scale surface structure patterns of birds’ flight feather, a novel thermal-protection enhancement technique is proposed for a turbine endwall with cavity air leaked from an upstream realistic rim seal. The potential application of the micro-scale surface structures is demonstrated by using well-validated numerical simulations. Overall film cooling performance and total thermodynamic loss are examined to comprehensively evaluate the feasibility of the micro-scale surface patterns for three cavity air flowrates. Detailed thermal fields and flow structures involved in the coolant–mainstream interactions are also presented to reveal physical flow mechanisms behind. Comparisons with a smooth endwall and previously used ribbed endwall cases show that applying the micro-scale structures onto the turbine endwall generates a substantial increase in cooling effectiveness with a very slight aerodynamic penalty. Despite of enhanced heat transfer levels, the resulting heat loads into the endwall are significantly reduced. Detailed thermal fields and flow structures revealed that the addition of the micro-scale structures induces counter-rotating vorticities that mitigate the passage vortex and hence prevents the accumulation of coolant near the suction side. Besides, it is also noted that the grooves of the micro-scale structures have a transport effect that can convey the coolant to the far downstream areas of the endwall, and thus considerably increase the coverage in the chordwise direction. The findings of this work provide an evidence that the bio-inspired, micro-scale surface structures could be an alternative for enhancing turbine endwall cooling performance.

Volumetric Velocimetry Measurements of Purge–Mainstream Interaction in a One-Stage Turbine

Abstract Air-cooled gas turbines employ bleed air from the compressor to cool vulnerable components in the turbine. The cooling flow, commonly known as purge air, is introduced at low radius, before exiting through the rim-seal at the periphery of the turbine discs. The purge flow interacts with the mainstream gas path, creating an unsteady and complex flowfield. Of particular interest to the designer is the effect of purge on the secondary-flow structures within the blade passage, the extent of which directly affects the aerodynamic loss in the stage. This paper presents a combined experimental and computational fluid dynamics (CFD) investigation into the effect of purge flow on the secondary flows in the blade passage of an optically accessible one-stage turbine rig. The experimental campaign was conducted using volumetric velocimetry (VV) measurements to assess the three-dimensional inter-blade velocity field; the complementary CFD campaign was carried out using unsteady Reynolds-averaged Navier–Stokes (URANS) computations. The implementation of VV within a rotating environment is a world first and offers an unparalleled level of experimental detail. The baseline flow-field, in the absence of purge flow, demonstrated a classical secondary flow-field: the rollup of a horseshoe vortex, with subsequent downstream convection of a pressure-side and suction-side leg, the former transitioning in to the passage vortex. The introduction of purge, at 1.7% of the mainstream flowrate, was shown to modify the secondary flow-field by enhancing the passage vortex, in both strength and span-wise migration. The computational predictions were in agreement with the enhancement revealed by the experiments.