Endwall Flow Research Articles

Endwall film-cooling in a cascade with filleted vane, upstream discrete slots, and holes

ABSTRACT In vane cascades, large filleted vanes influence aerodynamics, while upstream endwall film-cooling arrangement impacts aero-thermal performance. This study explores experimental investigations on endwall flows and film-cooling effectiveness in a filleted vane cascade using discrete slots and holes. Two fillet profiles are examined covering the leading-edge area to either the throat or 62% of the axial chord. Film-cooling employs two slots and two rows of holes in the leading-edge region with inlet blowing ratios ranging from 1.0 to 2.8. Results show variations in total pressure loss and Nusselt number between filleted and un-filleted vanes, highlighting optimal film-cooling effectiveness with the smaller fillet profiles.

A quantitative three-dimensional inverse design rule based on cross-flow control of profiled end wall

Profiled end wall provides a novel and effective solution for end wall flow control in turbines and compressors. However, the application of profiled end walls in compressors still lacks quantitative design rules. This paper presents a quantitative design rule for end wall profiling based on a three-dimensional inverse method and numerically validates it on a highly loaded compressor cascade. The current inverse method can solve the corresponding end wall shape from a given end wall pressure distribution, but the determination of the end wall pressure distribution heavily relies on empirical knowledge. This study establishes a model between streamline curvature and cross-passage pressure gradient (CPG) through the circumferential equilibrium equation in the S1 stream surface, thereby providing a quantitative basis for determining the end wall pressure distribution. The quantitative design rule proposed in this paper is expressed as follows: at the axial position where the separation begins on the suction surface (SS), within the range of 0.1-0.2 pitch away from the SS, the end wall boundary layer fluid with a higher velocity than the corner region average velocity should possess the same streamline curvature as the fluid within the viscous sublayer. The inverse-designed profiled end wall using the quantitative design rule enhances the local cross-flow near the SS by imposing a stronger CPG, thus encouraging the end wall boundary layer fluid with relatively higher momentum to arrive at the SS earlier and enhancing the radial migration on the SS. Consequently, the intensified cross-flow entrains relatively higher momentum into the corner region, while the enhanced radial migration drives the low-momentum fluid away from the corner region towards the midspan. Finally, the inverse-designed profiled end wall reduces the half-span mass-flow-weighted average loss by 4.3%.

A three-dimensional inverse method based on moving no-slip boundary for profiled end wall design

How to utilize existing flow control mechanisms to make profiled end wall design more flexible, efficient, and physical is a meaningful challenge. This study presents a three-dimensional inverse method for profiled end wall design to achieve the application of flow control mechanisms. The predetermined pressure distribution on the end wall is reached by modifying the end wall geometry during flow field calculation. A motion velocity model is derived from the normal momentum equation of the moving no-slip boundary to modify the end wall geometry. A Reynolds-Averaged Navier-Stokes (RANS) solver based on the Semi-Implicit Method for Pressure Linked Equations (SIMPLE) algorithm is adopted to simulate the flow field. Based on the mechanism understanding obtained through numerical optimization results, this study adopts the inverse method to redesign an optimized end wall in a compressor cascade. The results indicate that the redesigned end wall exhibits better loss reduction, reducing the overall total pressure loss by 5.5%, whereas the optimized end wall reduces it by 3%. The inverse method allows the imposition of desired influences on the end wall flow without constructing a database, making it highly flexible, efficient, and physical.

Sensitivity of the endwall flow in a linear vane cascade to blade fillet geometry

Based on the blade chord and inlet velocity, the current computational study uses a linear vane cascade with a large filleted blade-endwall junction with a 2.01 x 105 Reynolds number. Three fillets with related profiles are explored. To evaluate the upshots of geometric differences in a fillet attached to the endwall flow-field, the height and endwall-width of the fillets are changed. The RANS k-ω turbulent model is used in the computations, and the results are compared to experimental results from a similar cascade without the fillet. The computed results of the secondary flow-field in the endwall region along the cascade are compared for baseline (no fillet) and filleted passages. As a result of diminished leading-edge and passage vortices, the fillets lower pitchwise pressure gradients, flow separation, axial vorticity, and overall pressure losses when compared to the baseline. The pros of fillets on endwall secondary flows are however unaffected by fillet's geometric changes.

Unsteady Interaction Between Purge Flow and Secondary Flow in High-Lift Low Pressure Turbine

Abstract The coupling effect between the complex vortex system and the rim purge flow in the endwall region of a high-lift low pressure turbine (LPT) will significantly influence the evolution of secondary flow and the corresponding losses. This paper focused on the unsteady interaction mechanism between the rim purge flow and the secondary flow inside the high-lift LPT under the periodic wake passing. The large eddy simulation (LES) method was used to reveal the influence mechanism of rim purge flow, rotor-stator cavity interaction and unsteady wakes on the secondary flow. Detailed experimental measurement was carried out for the flow field of high-lift LPT under the influences of static purge flow. The results showed that the rim purge flow significantly increased the overturning and underturning downstream of the endwall region, resulting in aggravating secondary loss. The rotating-rim increased the difference of the circumferential velocity between the mainstream and the purge flow, aggravating the Kelvin-Helmholtz (K-H) instability, inducing the K-H vortex structure with stronger energy amplitude, and further increasing the endwall flow loss. The incoming wakes weakened the energy amplitude of the K-H vortex at the rim purge outlet, thinned the thickness of the low-energy fluid at the blade leading edge. Additionally, the interaction between the wakes and the secondary vortices further suppressed the development of the secondary flow. Nevertheless, the incoming wakes caused additional mixing losses and made a negative impact on the overall aerodynamic performance of the high-lift LPT.

An Aerodynamic Investigation of the Last-Stage Turbine in an Upgraded Gas Turbine

Abstract Due to the lengthy certification process for newly designed turbine blades, product upgrading of industrial gas turbine units is often performed solely on compressor and combustor. Since their inlet conditions are significantly changed, the entire four-stage turbine operates far away from its original design point, leading to decreased efficiency, and increased flutter risk. This investigation first performs numerical simulations to study the flow field change of the last-stage turbine in a gas turbine before and after product upgrading. To reduce the load of the last-stage turbine without reducing the power output of the whole turbine, the enthalpy drops of turbine are reallocated to the front three stages. After modifying the blade profile based on S1 stream surface analysis, a CFD simulation is carried out on the modified three-dimensional blade passage. It is shown that the modified blade design greatly reduces the Mach number at the tip outlet of the last-stage blade, thus possibly reducing flutter risk and improving the aerodynamic efficiency of the turbine. This paper also attempts to redesign the blade geometry by different radial blade stacking of both forward sweep and backward sweep. It is found that the backward-swept blade modification can effectively reduce the endwall flow loss. This work presents the improvements of the aerodynamic efficiency of last-stage through a series of improvement methods and provides a reference for future detailed optimization of this last-stage turbine.

Orifice Jet Curvature and Its Interaction With a Row of Short Pin-Fins

Abstract The fluidic behaviors and interactions present in a setup utilizing perforated blockages and downstream pin-fin are elucidated: specifically, the role of jet curvature in flow interactions and the interaction of end-wall flows around the pin-fins subjected to impinging jet. To this end, a combination of particle image velocimetry-based measurements and oil-dye surface flow visualization technique is utilized. Two different configurations of jet flows, with or without flanking jets, and three different distances between the pin-fins and blockages were considered. Jet curvature can result in flow inclinations of as much as 45 deg–60 deg interacting with the pin-fins and the majority of the jet flow is skewed toward the inner side of the pin-fin. At the end-wall, the interaction with nearby structures, being blockages or pin-fins, alters the shape and profile of the flows. This alteration is in contrast to those around conventional pin-fin arrays which remain largely unaltered in the presence of nearby pin-fins.

Flow mechanism of a highly loaded low pressure turbine cascade with integrated-optimized end wall contouring and root lean

The end wall loss of modern highly loaded low pressure turbine (LPT) has been greatly increased, due to the enhanced secondary flow loss and boundary layer separation loss. Thus, it is of great significance to develop effective flow control strategies to improve the end wall flow condition and aerodynamic performance of modern LPT. This research carried out a numerical investigation on the coupled flow control strategy, which combined non-axisymmetric end wall contouring (NEC) and root tangential lean (RTL), based on a highly loaded LPT cascade (Zweifel = 1.59). Meanwhile, the optimization process was used to get the optimal design parameters of the coupled method NEC&RTL. The results indicate that the optimal coupled configuration can reduce the total pressure loss coefficient by 12.68% and the non-dimensional secondary kinetic energy by 23.91%. Compared with the reference cascade without modification, the coupled method is found to improve the end wall flow conditions: the passage vortex is weakened both in size and strength, mainly attributed to the smaller cross-passage pressure gradient resulting from NEC; the closed separation bubble near end wall and the three-dimensional separation flow before trailing edge are eliminated, due to the great downward pressure gradient near end wall resulting from RTL; and the counter vortex is eliminated and the slender back flow is weakened under the additional coupling flow control effect of NEC&RTL. Therefore, the coupled flow control method can not only highlight the advantages of the independent methods, but also induce external flow control superiorities, demonstrating the application prospect of the coupled flow control strategy on the highly loaded LPT.

The Application of a Laser-Printed Miniature Five-Hole Probe in the End-Wall Flow Measurement of a Multistage Axial Compressor

To make measurement of end-wall flow between blade rows in a compact multistage configuration possible, a miniature L-shaped five-hole probe was employed in this paper. This compact tip structure, realized by laser-printing instead of the conventional machining technique, reduces the blockage effect of this intrusive measurement on the flow and ensures high spatial resolution. The zonal method is introduced to extend the usable flow angle range up to 60 degrees. A local least-squares interpolation technique is utilized to acquire flow angle and static/total pressure. In order to improve accuracy for the points located at the sector boundary, the overlap region method is included in the interpolation. Additional test data indicate that the maximum error in flow angle is nearly within 1 degree, and the maximum errors of total pressure and static pressure are 0.56% and 1.9% respectively. The application in a low-speed multistage axial compressor indicates that the zonal method can decrease the number of points exceeding the measurable flow range and is of great significance for the end-wall flow measurement, especially for the near-stall condition. Compared with the traditional method, the proportion of available data for the near-stall state measurement was increased by 18% by using the zonal method.

Synergy principle and its application of endwall loss analyses in the turbine stator

The endwall flow significantly impacts the turbine performance behavior, and it is necessary to investigate the development of endwall secondary vortices and relevant losses. Inspired by the field synergy principle, the synergy between the velocity and the pressure gradient established by the three-dimensional mechanical energy conservation equation is innovatively applied to the endwall loss analysis in the stator of an axial-inflow turbine. In terms of the synergy equation, the loss is not only related to the viscous dissipation, but also the included angle (or the synergy angle) between the velocity vector and the pressure gradient vector. The physical content of the synergy angle suggests that the larger synergy angle is (i.e., the worse synergy), the higher losses should be. This conclusion has been validated by present numerical results, and an apparent positive correlation between the synergy angle and the losses could be perceived under time-averaged and transient conditions. The worse synergy could be observed at the passage rear part and the wake, where complex passage vortices and local separation exist. In these regions, the local velocity vector is not aligned with the bulk pressure gradient of the mainstream. Hence, the synergy angle has a marked rise, corresponding to the local high losses.

Turbine Blade Endwall Heat Transfer and Film Cooling Performance With Multigap Jets and Film Holes

Abstract Due to manufacturing, assembly, and operational conditions, gaps are commonly presented in the turbine components, such as platform blade-to-blade gap (slashface gap) and rim cavity platform gap (slot gap). The coolant ejected through these gaps not only limits hot mainstream gas ingestion into the disk cavity but also has the potential to provide endwall film cooling coverage. Nevertheless, the interaction of these gap flows, predesigned film hole jets, and endwall secondary flows, significantly impacts endwall heat transfer and film cooling performance and leads to a more complicated flow field near endwall. To further understand the endwall flow physics behavior in this complicated flow field, the combined effects of film hole jets and multigap leakages (slashface jet and slot jet) were experimentally studied in a transient wind tunnel with a six-blade linear cascade (nonrotating). The endwall heat transfer was measured and recorded by the IR technique at inlet average turbulence intensity (Tu) of 7.5% and an exit Mach number (Maex) of 0.4. In addition, detailed numerical predictions were also performed to discuss the flow physics near endwall and the multigap leakages behavior. Results indicated that the slashface jet is an essential contributor to endwall film cooling performance in this endwall configuration. The slashface jet delays the development of the film hole jets, leading to an enhancement of film cooling performance downstream of the film holes. Increasing the mass flow ratio of the slashface jet (MFRslashface) can prevent the high-temperature mainstream ingestion, reduce the thermal failure risks, and increase the peak of film cooling effectiveness downstream of the endwall (167% at MFRslashface = 0.5%). Due to the limitation and hindrance impacts of the disk vortex (DV), the cavity vortex (CV), and the suction side leg of the horse-shoe vortex (HV,s), the slot flow is confined to a small region downstream of the slot gap.

Inception of Casing Wall Flow Instabilities and Downstream Flow Evolution in Subsonic Axial Compressor Rotors At High Rotor Stagger Angles

Two critical aspects of the casing wall flow structure in a subsonic axial flow compressor are studied in this paper: flow mechanism leading to stalling at different stagger angle configurations; and the effect of tip leakage flows on these flow phenomena. In some of the previous works, casing wall flow instabilities similar to the ones discussed in the present paper were reported, typically for high stagger angle rotor designs. However under what conditions flow tends to take such shapes were not discussed in detail. This work attempts to address this through a detailed numerical study on unswept and forward swept rotors, for different stagger angle settings, and for varied levels of rotor tip gaps. Three candidate rotors (unswept, 20° true swept, 20° axially swept) are studied with three tip clearances (0.0%, 0.7%, and 2.7% of the blade chord) for three stagger angle changes (0°, +3° and +5°). Stagger angle is found to have significant effect on suction surface flow separations. Tip clearance is observed to be critical on the inception of casing wall flow instabilities, especially at higher stagger angles. Hub endwall flow structure and the downstream flow evolution are also discussed in the paper.

Reduction in endwall flow losses in turbine blade with the use of sweep in leading edge of the blade

Reduction in endwall flow losses in turbine blade with the use of sweep in leading edge of the blade

Upstream Jet Cooling and Dual Cavity Slashface Leakage Cooling on a Transonic Nozzle Guide Vane Endwall

Abstract This paper presents an experimental and computational study on implementing a dual cavity slashface cooling scheme on the thermal performance for the first stage nozzle guide vane with an axisymmetric, converging endwall. An upstream dual-row, staggered cylindrical hole cooling scheme provided purged coolant. The study was conducted under representative engine conditions of Maexit = 0.85 and Reexit,Cax = 1.5 × 106. Data were collected and analyzed using infrared thermography to map the endwall heat transfer performance throughout the passage. A flow visualization study was employed to gather qualitative insights into the endwall flow field. In addition, a complimentary computational fluid dynamics study was carried out to understand the endwall flow ingestion and egress behavior near the slashface. Results indicate that the dual-plenum slashface scheme leads to a unique ingestion-egression-ingestion-egression pattern. The ingestion and egression suppress the upstream coolant attachment at mid-passage near the suction side but create a favorable coolant coverage downstream of the slashface tail-end, where the neat heat flux reduction is 0.4–0.5. The study also compared the dual-plenum design with the single-plenum design and concluded that the dual-plenum slashface is superior in film cooling performance.

Experimental study on the outlet flow field and cooling performance of vane-shaped pre-swirl nozzles in gas turbine engines

The pre-swirl nozzles are a crucial part of gas turbine's cooling pre-swirl system, because they directly affect the cooling performance of the system and the life of the turbine blades. Vane-shaped pre-swirl nozzles, including cascade vane nozzles (VN) and vane-shaped hole nozzles (VSH) proposed in recent years, are reported to have higher system temperature drop potential. In this paper, the pre-swirl performance as well as the cooling effect of these two types of nozzles were studied and compared experimentally. A five-hole probe was used to measure the flow field at nozzle exits under the outlet conditions of Mach number 0.2 and 0.3, and the comparisons of global deviation angle and pre-swirl effectiveness were carried out. The variations of pre-swirl effectiveness (η) with aerodynamic parameters, such as pressure ratio and Reynolds number in the range of 3 × 104 to 8 × 105 were also obtained with the help from numerical simulations. The cooling performance of two nozzles were further tested by a system test rig with rotating parts downstream of the two nozzles. Total pressure and outlet flow angle at nozzle exit illustrate that VN have obvious boundary layer separation from the suction surface, and the secondary endwall flow loss is high due to their narrow flow passages. As a result, when Mach number is 0.2, the deviation angle for VN is about 3° and the pre-swirl effectiveness is only 0.77. Well-designed VSH have improved flow passages with appropriate height to pitch ratio and a smaller trailing edge, and they suppress boundary layer separation and secondary flow development. This contributes to a significantly lower deviation angle near 0°, and an increased pre-swirl effectiveness of 0.86. In addition, the pre-swirl effectiveness of these two vane shaped nozzles will first increase with Reynolds number and then tend to a constant at Reynolds number about 4 × 105 and beyond. The maximum effectiveness of VN is 0.85 and that of VSH is about 0.94 and is 10.6% higher. Meanwhile, the measured temperature drop effectiveness of the pre-swirl system with VSH reaches 0.49, achieving an 8.9% improvement over the system with VN at 0.45. These results are consistent with the theoretical linear relationship between system temperature drop and nozzle pre-swirl effectiveness. The investigations in this paper also conclude that the pre-swirl effectiveness of nozzles is a more accurate parameter in classifying the nozzle performance and system cooling effect than traditionally used discharge coefficient.

Toward the Utilization of 3D Blading in the Cantilevered Stator from Highly Loaded Compressors

Three-dimensional blading is an efficient technique in compressor aerodynamic design, and its function mechanism in the cantilevered stator needs to be addressed. This paper focuses on the sweep and dihedral in the cantilevered stator and seeks to expose their effects through detailed flow field analysis. Results show that the forward sweep could alleviate the corner flow separation by preventing the accumulation of the secondary flow toward the corner region, resulting in stronger flow separation at the blade trailing edge; in summary, forward sweep with appropriate parameters could increase static pressure rise by 14.3%. The positive dihedral will carry the endwall flow to the upper-span sections, thereby reducing blade corner separation; hence, as much as 23.5% improvement in static pressure rise could be obtained with the appropriate dihedral. Moreover, the combination of a relatively large sweep height and a moderate sweep angle with a low dihedral height and a moderate sweep angle provides optimum aerodynamic performance; the static pressure rise coefficient sees an increment of 25.5% at the near stall point. An experiment is then performed to further validate the theory, which shows a 2% improvement in efficiency of 3D blading at small mass flow rates. However, the secondary leakage should be given attention at high mass flow coefficients, while the corner separation needs further elimination at small mass flow rates.

Numerical Investigations on Effect of Inflow Parameters on Development of Secondary Flow Field for Linear Low-Pressure Turbine Cascade

Abstract Endwall flows contribute the most crucial role in loss generation for axial flow turbine and compressor blades. These losses lead to modifying the blade loading and overall performance in terms of stable operating range. The present study aimed to determine the endwall flow streams in a low-speed low-pressure linear turbine cascade vane using a numerical approach. The study includes two sections. The first section includes an attempt to understand different secondary flow streams available at the endwall. The location of the horseshoe vortex and subsequent vortex patterns are identified in the section. The selection of a suitable turbulence model among shear stress transport (SST) k–ω and SST γ–θ to identify endwall flow streams is studied prior to the section. The steady-state numerical study is performed using Reynolds Averaged Navier–Stokes equations closed by the SST γ–θ turbulence model. The computational results are validated with experimental results available in the literature and are found to be in good agreement. The study is extended for different inflow conditions in a later section. The second section includes the effect of flow incidence and turbulence intensity on the endwall secondary flow field. Four different inflow incidences are considered for the study. The inlet turbulence intensities are varied by 1% and 10% for each case. The results revealed different secondary flow patterns at an endwall and found the change in behavior with inflow conditions. SST γ–θ turbulence model with lower turbulence intensity is more suitable to identify such flow behavior.

Comparative numerical investigation of wake effect on low-pressure turbine endwall flow

In low-pressure turbine linear cascade experiments, the upstream stator vanes are often modeled by moving bar wake generators. Different from the majority of the earlier wake effect research which is concerned with less aggressively loaded airfoils, here front-loaded high-lift airfoils are considered for both stator and rotor. Implicit large-eddy simulations of the flow through a 50% reaction stage with stator vanes and rotor blades as well as a linear cascade with bar wake generators and rotor blades were carried out and the wake effect on the flow through the rotor was investigated. The chord-based Reynolds number for the simulations was 100,000. For the chosen front-loaded high-lift airfoil the stator wake is dominated by pronounced endwall structures which are lacking for the bar wake generator. Instantaneous flow visualizations reveal a periodic suppression of the two-dimensional rotor blade suction-side separation for both configurations. With upstream stator vanes, the wake endwall structures appear to be chiefly responsible for the suppression of the rotor blade passage vortex and corner separation. For the bar wake generator, the wakes are wider and the wake effect is spread out over a larger portion of the wake passing period. The passage vortex and corner separation remain mostly intact.

Low Reynolds Number Effects on the Endwall Flow Field in a High-Lift Turbine Passage

Abstract Understanding the endwall flow phenomena surrounding low-pressure turbine blades is key to improving performance, as these flow features contribute significantly to loss generation at low Reynolds number cruise. It is well documented that a horseshoe vortex system forms at the junction of the endwall and turbine blade. The vortices develop and gain significant strength in the passage and contribute to total pressure losses. During low Reynolds number conditions, the flow through a low-pressure turbine passage can be greatly impacted by a number of factors, including Reynolds number and incoming turbulence. The focus of this paper is on significant changes to the endwall flow field observed in experimental measurements and an accompanying implicit large-eddy simulation of the flow through a linear cascade of high-lift front-loaded low-pressure turbine blades at low Reynolds number. Results show a significant effect on both the time-averaged endwall flow topology and the unsteady vortical flow characteristics when the Reynolds number based on inlet conditions was decreased to 30,000. Various techniques, such as spectral proper orthogonal decomposition, were used to analyze and compare both high-speed particle image velocimetry measurements and numerical results in order to extract the dominant structures and their unsteady behavior. The total pressure loss development through the passage was assessed in order to better understand how the observed changes in endwall flow structures contribute to the overall losses.

Vortex Unsteadiness in the Endwall Region of a High-Lift Low-Pressure Turbine Passage

Abstract Three-dimensional vortical structures within the endwall region of turbine passages directly affect the aerodynamic efficiency and heat transfer characteristics of the turbine. Interactions between the vortical endwall structures and the suction surface flow have been shown to be a significant source of loss generation through passages. One dominant vortex extends from the leading-edge junction region of the blade across the passage, where it interacts with the flow along the suction surface of the adjacent blade. In high-lift low-pressure turbine cascade passages, the vortical structure intermittently loses coherence and exhibits unsteady variations of strength and position as it extends across the passage. The present paper details the temporal behavior through high-speed measurements in a low-speed linear cascade of high-lift low-pressure turbine blades. Stereoscopic particle image velocimetry measurements in the passage are used to evaluate the unsteady behavior of the vortex. Space-time iso-surface plots of Q-criterion clearly show the evolution of the vortex over time. Analysis of the data reveals the various time scales of fluctuations in strength and position. Comparisons of the temporal fluctuations in the high-lift turbine passage are made with similar phenomena found in canonical junction flow papers in the literature. Key findings support the hypothesis that in-passage vortex unsteady characteristics near the endwall are influenced by leading-edge junction flow dynamics, and provide additional insight into the unsteady endwall flow physics that is necessary to further the development of endwall loss reduction techniques.